4" - \ NATIONAL AERONAUTICS AND SPACE ADMINISTRATION Technical Report No. 32-881 ,,--:-^- _,_._ la_A Project Report:IVICLI II it_l I Vial o ! _ _ -7 . _. Mission Operations GPO PRICE $ CFSTI PRICE(S) $ Hard copy (HC) , _.7_, _ Microfiche (MF) /_' _S ff 653 July 65 (ACCESSION NUMBER) ==o _0/ 35 947 ...... (TH 71 (CODE) :31 (CATEGORY) PASADENA, CALIFORNIA June 15, 1966 https://ntrs.nasa.gov/search.jsp?R=19660026657 2020-03-16T20:34:43+00:00Z NATIONAL AERONAUTICS AND SPACE ADMINISTRATION T^^_-;_°/ Report No. 32-881 Mariner Mars 1964 Project Report: Mission Operations JET PROPULSION LABORATORY CALIFORNIA INSTITUTE OF TECHNOLOGY PASADENA, CALIFORNIA June 15, 1966 Copyright © 1966 Jet Propulsion Laboratory California Institute of Technology Prepared Under Contract No. NAS 7-100 National Aeronautics & Space Administration JL JPL TECHNICAL REPORT NO. 32-881 PREFACE On November 28, 1964, Mariner IV was successfully launched on a trajectory that would take it within 150,000 miles of Mars. On December 5, 1964, the spacecraft performed a successful midcourse maneuver, altering its flight path so that it would pass within 6118 miles of the planet. The flight took approximately 7½ months, during which time a great _1_1 of sei,_-_ff_o information was _atheredIdltK_ClkJt _A_ concerning the environment of both near-Earth and interplanetary space. And then, on July 14, 1965, Mariner IV photographed the sur- face of Mars, and telemetered to Earth the most advanced scientific and technical data regarding the planet yet recorded. An earlier volume of the series reports on Mariner III. This volume, which deals with Mariner IV, only, describes various operations of the Project during the flight from the time of midcourse maneuver to the end of the mission on October 1, 1965. It includes an account of the trajectory and orbit determination analyses, a description of the Space Flight Operations System, a summary of the post-launch testing results, and a brief account of the final configuration and disposition of the MC-4 and PTM spacecraft at the end of the mission. ( Since the gener- ation of material for this report, the additional Mariner IV and Mariner Venus 67 Projects were authorized by NASA, and these two spacecraft are being used in the follow-on projects. ) The Mariner Mars 1964 mission required the use of a great many new techniques in designing, building, and operating unmanned space- craft. The success of Mariner IV has made these techniques significant, particularly in the light of future space programs. It is hoped that the operations and results documented in this volume will be useful refer- ence to those who are planning fu__ W.H. ring, Director Jet Pro Assistant Laboratory , Project Manager Mariner Mars 1964 III JPL TECHNICAL REPORT NO. 32-881 ELEMENTS OF THE MARINER MARS 1964 PROJECT REPORT The Mariner Mars 1964 Project Report consists of the following volumes: Mariner Mars 1964 Project Report, Mission and Spacecraft Develop- ment, Volume I: From Project Inception Through Midcourse Ma- neuver (TR 82-740, Vol. I) Mariner Mars 1964 Project Report, Mission and Spacecraft Develop- ment, Volume II: Appendixes (TR 82-740, Vol. II) Mariner Mars 1964 Project Report, Mission Operations (TR 8.2-881) Mariner Mars 1964 Project Report, Spacecraft Performance and Analysis ( TR 82-882) Mariner Mars 1964 Project Report, Scientific Experiments (TR 82-888) Mariner Mars 1964 Project Report, Television Experiment, Part I: Investigators" Report (TR 82-884, Part I) Mariner Mars 1964 Project Report, Television Experiment, Part II: Picture Element Matrices (TR 82-884, Part II) Tracking and Data Acquisition Support for Mariner Mars 1964, Volume II: Cruise to Post-Encounter Phase (TM 88-289, Vol. II) IV JPI- TECHNICAL REPORT NO. 32-881 CONTENTS I. Project Organization and Management .......... 1 A. Structure ..................... 1 B. Schedules .................... 1 C. Reviews ..................... 3 D. Problem/Failure Reports ............... 3 E. Meetings ..................... 3 F. Workshops .................... 3 II. Trajectory and Orbit Determination ............ 4 A. Trajectory ..................... 4 1. Launch Phase .................. 4 2. Near-Earth Phase ................. 4 3. Midcourse Maneuver Phase ............. 6 4. Cruise Phase ................... 8 5. Encounter Phase ................. 10 6. Post-Encounter Phase ............... 12 B. Orbit Determination, Results .............. 13 1. Tracking Data: Validity and Usage ........... 15 2. Mariner IV Post-Flight-Encounter Orbit Tracking from E - 5 Days to E + 5 Days ............ 16 3. Evaluation of Combined Pre-Midcourse Orbit and Maneuver Errors ................ 17 4. Cruise Orbit Determination Errors ........... 18 5. Near-Encounter Orbit Errors (Epoch 0_ 10 July 1965; Tracking Data from 0h 10 July) ............ 19 6. Conclusions ................... 19 III. Space Flight Operations ................ 22 A. Data Recovery and Processing ............. 22 1. Science and Engineering Data ............. 22 2. Tracking Data .................. 22 3. Coverage .................... 22 4. Ground Telemetry Subsystem Performance ........ 23 5. Picture Data ................... 23 6. Occultation Phase ................. 23 7. Mariner Master Data Library ............. 24 B. Space Flight Operations Performance ........... 24 1. Investigation and Correction of Problems ......... 27 2. Three-Shift Coverage, 6 July to 6 August 1965 ....... 28 3. Support Services Rendered ............. 28 V JPL TECHNICAL REPORT NO. 32-881 CONTENTS (Cont'd) C. Encounter Readiness Tests ............... 31 1. Roll Increment Tests ................ 31 2. Scan Platform Positioning .............. 33 3. Procedure Verification Tests ............. 33 4. Backup Mode Test ................. 33 5. Tests from Flight Path Analysis Area .......... 33 6. Operational Readiness Tests ............. 34 7. Command-Loop Loekup Exercises ........... 34 IV. Post-Launch Testing .................. 35 A. Life Test Program .................. 35 1. Significant Findings ................ 37 2. Summary of Significant Interruptions .......... 39 3. Recommendations for Future Programs ......... 40 4. Conclusions ................... 41 B. Electron Radiation Tests ............... 41 1. Environmental Estimates .............. 42 2. Test Activity ................... 43 3. Results .................... 44 4. Summary .................... 46 C. Electromagnetic Interference Test Activities ......... 46 1. Spectrum Signature Test of PTM ............ 47 2. RF Simulation Test with Peak RF Power Densities ...... 52 3. S-Band RF Noise Measurements on PTM in Spacecraft Assembly Facility ............. 52 D. Magnetic Stability and Demagnetization Tests ........ 56 1. Solar Panel Magnetic Stability Tests .......... 56 2. Proof Test Model Bus Demagnetization Study ....... 57 E. Proof Test Model, Usage and Disposition .......... 59 1. From 1 to 8 December 1964 ............. 59 2. From 9 December 1964 to 11 February 1965 ........ 60 3. From 12 February to 14 July 1965 ........... 63 4. From 16 July to 23 September 1965 ........... 67 5. Spacecraft Configuration for Storage .......... 69 F. MC-4 Spacecraft Operations and Disposition ........ 70 1. From 30 September 1964 to 9 April 1965 ......... 70 2. From 9 April to 15 June 1965 (at JPL) .......... 72 V. End-of-Mission Planning ................ 80 A. Proposal for 1967 Reacquisition ............. 80 1. Introduction ................... 80 2. Objectives .................... 81 3. Alternate Plans .................. 83 VI JPL TECHNICAL REPORT NO. 32-88! CONTENTS (Cont'd) B. Spacecraft Final Conditioning ............. 84 References ....................... 85 Appendixes A. P List ...................... 86 B. Schedule Package .................. 88 C. Summary of Reference Documents ............ 100 D. Encounter Sequence of Events ............. 127 TABLES 1. Major review sessions .................. 3 2. Geocentric characteristics of Mariner IV trajectory ......... 6 3. Heliocentric orbital elements of Mariner IV trajectory ........ 8 4. Areocentric orbital elements of Mariner IV trajectory ........ 9 5. Computer printout distribution ................ 29 6. SFOF Mariner documents .................. 30 7. Encounter readiness tests .................. 31 8. Command and occultation tests ............... 33 9. Command-loop Iockup exercises ............... 34 10. Summary of life test status ................. 35 11. Subsystem failure modes ................. 44 12. RF simulation test parameters ............... 53 13. Mariner C PTM residual field at Magnetometer .......... 59 14. Test sequence, 12 February to 14 July 1965 ........... 64 15. Test sequence, 15 July to 23 September 1965 .......... 67 16. PTM configuration and nominal characteristics at time of storage 69 17. Operations and tests of MC-4 at AFETR ............ 71 18. Tests at JPL prior to storage ................ 75 19. SAF initiated P/FRs against MC-4 spacecraft .......... 77 20. MC-4 configuration and nominal characteristics at time of storage 78 21. Canopus sensor position and command sequence for 1966 and 1967 81 VII JPL TECHNICAL REPORT NO. 32-881 i FIGURES 1. Mariner Project organization chart .............. 2 2. Ascent trajectory profile .................. 4 3. Sequence of events to Canopus acquisition ........... 5 4. Earth track of Mariner IV trajectory .............. 7 5. Heliocentric plan view of Mariner IV trajectory during first 35 days of flight .................... 8 6. General relationship of Earth, Mars, and Mariner IV during cruise to encounter ................... 9 7. Geocentric radius vs time from launch ............. 9 8. Geocentric speed vs time from launch ............. 10 9. Heliocentric radius vs time from launch ............ 10 10. Heliocentric speed vs time from launch ............. 10 11. Areocentric radius vs time from launch ............. 10 12. Earth cone angle vs time from launch ............. 10 13. Earth clock angle vs time from launch ............. 10 14. Mariner IV and Mars orbits near encounter ........... 11 15. Mariner IV encounter of Mars as seen from inside Mars orbit ..... 11 16. Heliocentric plan view of Mariner IV pre- and post-encounter orbits . 12 17. Spacecraft celestial latitude vs calendar date .......... 12 18. Geocentric radius vs calendar date .............. 13 19. Heliocentric radius vs calendar date .............. 13 20. Areocentric radius vs calendar date .............. 14 21. Earth cone angle vs calendar date .............. 14 22. Earth clock angle vs calendar date .............. 15 23. Mariner IV errors in predicting B'Tc ............. 18 24. Mariner IV errors in predicting B" Rc ............. 18 25. Mariner IV errors in predicting time of closest approach ....... 18 26. Errors in predicting encounter quantities vs time, Mariner IV ..... 19 27. Mass of Mars solutions .................. 20 28. AU solutions from Venus bounce and Mariner IV data ........ 21 29. Operations area .................... 25 30. Spacecraft performance analysis area ............. 25 31. Space Science Analysis area, first floor ............. 25 VIII JPL TECHNICAL REPORT NO. 32-881 FIGURES (Cont'd) 32. Flight Path Analysis area, first floor .............. 26 33. Mariner Mission Support area ................ 26 34. Track area ...................... 27 35. M_r,n_r -=n_'_,J.-..terG.mund Communications ........... 32 36. Electron radiation test setup ................ 42 37. Electron flux distribution in the area of the test flux ........ 43 38. Radiation thresholds of malfunction for various subsystems in test 44 39. Performance of the roll error signal under radiation with electronics at 2.25 Mev ................ 45 40. Apparent star intensity far no star input vs radiation flux rate .... 45 41. Block diagram of test configuration .............. 47 42. Spectrum signature, dipole probe .............. 43. Spectrum signature, rod antenna .............. 49 44. Spectrum signature, horizontal dipole ............. 50 45. Spectrum signature test results, spacecraft on vertical and horizontal polarization .................. 51 46. Simulation equipment .................. 52 47. Modulation source for Atlas TM-1 and Agena TM sources ...... 52 48. Noise receiver configuration for S-band noise measurements ..... 49. Mariner X-Y plane demagnetization facility with Earth's field bucking Helmholtz coils ................ 58 50. Demagnetization of PTM spacecraft in Z axis with Earth's field bucking Helmholtz coils ................ 58 51. Mariner PTM test and operations summary ........... 61 52. MC-4 test history, 25 November 1964 to 24 May 1965 ....... 73 53. MC-4 PIPS positioning .................. 76 54. Interplanetary missions coverage .............. 82 55. Solar activity and launch date ................ 82 IX JET PROPUI_ION LABORATORY INTEROFFICE MEMO 19 June 1963 TO: FROM: SUBJECT: All Concerned J. N. James Significance of 19 June 1963 Today, the 19th of June, the USSE spacecraf_Msrs _msde its encounter with Mars - dead 8s a doornail. You are receiving an issue of the memorandum because you are one of the individuals who by your daily actions and efforts can make Mariner C the first space- craft to take measurements on the planet Mars. It's as tough a job as we could pick. The Soviets have made st least seven launches to Venus and Mars, none of which have succeeded. You have tried twice - to Venus - and succeeded with Mariner II. On the basis of those statistics you are better than they are. ....... n...._ _=_ of statistics we should look a_. .An analysis* made for ARPA by Arinc Research, Inc. shows that of sixteen satellites that had been injected, they demonstrated a 95% probability of living only 2000 hours. Mariner II ceased to operate after 129 days - s little over 3000 hours. The Mars I, if we can believe the Soviet news releases, was launched 1 November and failed 21 March, giving it a life of about 3400hours. We must launch t_ Mariner C spacecraft in 1964 which will live at least---_OOhours. To accomplish this depends on each of you as an individual - your initiative - your craftsmanship - your ingenuity - your precision - your making each decision in the direction of success - your conscientious attitude towards the system as a whole and not just your part of the system slone- your meetin6 of schedules on time to permit us to face and resolve the unknowns quickly. I believe we have a first rate design in the Mariner C. We have plenty of talent assigned to all Project area_ and the schedule isn't too bad. So it is pretty much up to each of us to make every day count. The Soviets will launch again for Mars in 1964 but they will have some company. Thus you have another opportunity to demonstrate to the world that you excel in this type of venture by encountering Mars with two spacecraft in perfect operating condition. JNJ:pk *See Aviation Week, May 28, 1962 JPL TECHNICAL REPORT NO. 32-881 I. PROJECT ORGANIZATION AND MANAGEMENT A. Structure Responsibilities for key divisions of mission activities for the flight and encounter phases of Mariner IV were essentially the same as those discussed at length in the Mission and Spacecraft Development volume for the pre- lxajectory correction maneuver phase. Figure 1, which defines the specific, applicable, organizational structure for the period reported in this volume, reflects the greater emphasis on flight analysis and support activities and the lesser emphasis on the design and development work- except as it pertained to unusual equipment performance or non-standard telemetry results. One function not shown on the chart was the necessary one of public information, which drew a considerable amount of Project personnel's attention; this requirement grew in proportion to the suc- e,:ss of the mission. The management structure provided the base for vari- ous administrative checks and controls of the project; these took such forms as master planning and scheduling and follow-up, reviews, reports, and appropriate meet- ings and workshop sessions. In addition, the Project P List was continued throughout the mission to highlight major problems that could seriously affect the success of the mission. The post-launch P List is presented in Appendix A. B. Schedules The absolute irrevocability of flight events imposed even more stringent demands on the precision and thor- oughness of scheduling than did the pre-liftoff events. For example, the encounter sequence necessitated more exact time-table adherence than did countdown-obvi- ously, encounter could not slip one day, as did launch. That the schedule was not only designed but executed within proper tolerances is evidenced by the success of the mission. Details of steps and objectives, and a record of the manner in which they were met, are shown graph- ically in the schedule package in Appendix B. These charts present original schedules, completion dates, and rescheduling information. Many of the highlights of flight support are contained in these charts. Here again, as in the case of the organization chart for the Project, JPL TECHNICAL REPORT NO. 32-881 _ z d m gg>_ ._1 _ z __o._ >_-3_ ZI-- Z °" _--g. J !0 n" W (n 0 Z aJ g H_ w ,_1 c0u:- -- -wz m _w Rz >-S: z,..0 _o 0 Z c_m _--Z r I m ___L__. _j o z_c_ w -_ >. _ o_ -J _ u_Q: z I w I u _--_1 Z uol Nz _ o_ _ _j x i 3 ° u_ w 0 O_ tjj_ w _ O _J w woo ._1 _j _ (D -Ju_ _ _J w _ L._ O. I-- 00. _, 00. zuJ z _-_ o_ _ w ._..,- I =_ o_- I _ _- I LLI ,W _J W - 3 o 3 1 , m m _- t _.) O: J 2 __ u_ ,'r I _ c o _ g _ _ _ "-E W_Z _ z _ O. _ uJ T _ _ )Z .0 I-- 0 I-- _ _ _ o_ _ _ bJ Z _- 8 _ _o z _> Oo_ ._ "_ o 0 -_ _-- n." (/I u (.j 0") <_ _c_ .j : 0 _ o I- z _ _ o_ _ _ _ Z 0 ._1 ¢nn,- _ mm _" B 8 __ 2 ..o JPL TECHNICAL REPORT NO. 32-881 there is a strong resemblance between this group of charts and that included in the volume covering the earlier period of mission activity. C. Reviews One of the most effective techniques used in assuring that problems were receiving appropriate attention was in the use of continuing reviews-regular Quarterly Re- views of the Project by NASA I/cadqu_e_, and self review of selected areas of concern by the Project, itself. These latter were more frequent as time for encounter neared. Table 1 lists some of the major sessions. Table 1. Major review sessions Date, 1965 Subject Location 23 February 3 May 17 May 25 May 30 June 23 July 2 August 28 October Quarterly review Encounter planning Occultation planning Quarterly review Encounter preparation status Encounter review Quarterly review Quarterly review NASA HQ JPL JPL JPL JPL JPL JPL NASA HQ Encounter-readiness test critiques continued up to the day before fly-by. Finally, a post-encounter briefing was held to disseminate data to assure all Project participants' receiving the important information. D. Problem/Failure Reports During the subject period of the mission, problem/ failure reports (PFRs) continued to be made on flight performance anomalies and on misoperation observed during the extended testing of the spare, MC-4 space- craft, and of the proof test model (PTM), MC-1. Similar reports were prepared by the operations team against the Space Flight Operations Facility (SFOF); and when oc- casioned, malfunction reports were generated because of difficulties of the Deep Space Network (DSN). In response to these notices, the cognizant organization was required, as in the development phase, (1) to under- take a thorough and responsible diagnosis of the prob- lem, (2) to establish corrective and preventive measures, (3) to assure by supervisor signoffs that the best remedial steps had been taken, and (4) to follow up all problem areas by reviews. A summary of PFRs appears as an appendix in Vol- ume I of this Mariner Mars report. E. Meetings Unless superseded by other more important meetings, there were weekly meetings of Project Representatives. In addition, daily flight operatiol_s briefings were, held. From these latter sessions, in which were discussed unusual spacecraft or science activities, were originated the daily technical bulletins and TWX s to Headquarters. Shortly after the trajectory correction maneuver, the Space Flight Operation Director's team instituted a series of regular meetings devoted to encounter planning. Every facet of the encounter was probed and an optimum en- counter plan was evolved. Other groups meeting to solve particularly difficult tasks were those concerned with occultation and data processing operations. F. Workshops In the design, development, and testing of the Mariner IV spacecraft, use was made of certain tech- nologies that reached new limits. Since, in some cases, requirements could not be met by use of existing meth- ods, new techniques were evolved to meet the unique engineering demands of the mission. Steps were taken to disseminate information of these advanced technologies to aerospace affiliated personnel and to others who might benefit from its application. Project members were encouraged to contribute to, and participate in, professional conferences and symposiums and to submit papers for professional journal publication. Three areas of serious concern to the Mariner Project were the subjects of technical workshops hosted by the Jet Propulsion Laboratory, with attendance ranging from 50 to 150. The Magnetics Workshop (Ref. 1) was held 30 March to 1 April 1965; the Thermal Workshop met 23-24 June; and the High Voltage Workshop (Ref. 2) convened 18--20 August 1965. In addition to the above cited proceedings, a list of available documents pertaining to many aspects of Mariner IV history are listed in Appendix C. 3 JPL TECHNICAL REPORT NO. 32-881 II. TRAJECTORY AND ORBIT DETERMINATION A. Trajectory 1. Launch Phase The Mariner IV spacecraft was launched from Launch Complex 12 at the Air Force Eastern Test Range (AFETR), Cape Kennedy, Florida, on Saturday, 28 November 1964, using an Atlas D/Agena D launch vehicle. Launch occurred at 14:22:01.309 CMT, with an inertial launch azimuth of 90.5 deg east of north. After liftoff, the booster rolled to an azimuth of 91.4 deg and performed a programmed pitch maneuver until booster cutoff. During the sustainer and vernier stages, adjust- ments in vehicle attitude and engine cutoff times were commanded, as required, by the ground guidance com- puter, to adjust the altitude and velocity at Atlas vernier cutoff. After Atlas/Agena separation, there was a short coast period prior to the first ignition of the Agena engine. At a preset value of sensed velocity increase, the Agena engine was cut off. At this time, both the Agena and the attached spacecraft were coasting in a nearly circular parking orbit in a southeasterly direction at an altitude of 188 km and traveling at an inertial speed of 7.80 km/sec. After a parking orbit coast time of 32.25 min, deter- mined by the ground guidance computer and transmitted to the Agena during the Atlas vernier stage, a second ignition of the Agena occurred. Ninety-six seconds later, the Agena was cut off with the Agena-spacecraft com- bination in a nominal Earth-Mars transfer orbit. The launch phase ascent trajectory profile is illus- trated in Fig_ 2, and a sequence of events, from launch to Canopus acquisition, is shown in Fig. 3. 2. Near-Earth Phase Injection (second Agena cutoff) occurred at 15:04:27.7 GMT over the Indian Ocean at a geocentric latitude and longitude of -26.25 and 68.82 deg, respectively. At that time the Agena-spacecraft combination was at an altitude of 197.2 km and traveling at an inertial speed of 11.448 km/sec. The geocentric characteristics of the Agena-Mariner IV trajectory are listed in Table 2. At 1 min 23 sec after injection, the Agena combination entered the Earth's shadow. The Agena separated from the spacecraft 1 min, 18 sec later. At 15:17:34.8 GMT, Mariner IV left the shadow after a total shadow duration of 11 min, 44 sec. Within an hour after injection, the spacecraft was receding from the Earth in almost a AGENA D SECOND I GNITION D FIRST CUTOFF --AGENA D SECOND CUTOFF .... D/SPACECRAFT SEPARATION E c ,S :D I-- I- J 400 500 200 100 IOOO 2000 3000 VECO SECO -BECO FIRST IGNITION RANGE, nmi Fig. 2. Ascenttrajectoryprofile 4000 500C 6000 4 JPL TECHNICAL REPORT NO. 32-881 13 12 II EVENT I. LAUNCH 2. BOOSTER ENGINE CUTOFF AND JETTISON 3. SUSTAINER CUTOFF 4. VERNIER CUTOFF 5. SHROUD EJECTION 6. AG£'NA SEPARATION 7. AGEN.4 FIRST-BURN IGNITION 8. PARKING ORBIT INJECTION 9. AGENA SECOND-BURN IGNITION I0. SECOND AGENA CUTOFF (:INJECTION) I I. SPACECRAFT SEPARATION 12. SPACECRAFT SOLAR ACQUISITION 13. SPACECRAFT CANOPUS ACQUISITION 8, 9, I0 j13 \\ \\ 3,4 Rg. 3. Sequence of events to Canopus acquisition 5 JPL TECHNICAL REPORT NO. 32-881 Table 2. Geocentric characteristics of Mariner IV trajectory Characteristic Pre-encounter (injection) Pre-encounter (post-midcourse) Post-encounter Parameter Radius R, km Inertial speed V, km/sec Earth-fixed speed v, km/sec Geocentric latitude _, deg 6872.9574 11.206585 10.775777 --28.130141 Longitude 0, deg Right ascension O, deg Path angle of inertial velocity I', deg Azimuth of inertial velocity _, deg Path angle of Earth-flxed velocity y, deg Azimuth of Earth-fixed velocity a, deg Time of event T, GMT Hyperbolic orbital element Semimajor axis a, km Eccentricity e Inclination to Earth's equator i, deg Longitude of ascending node _, deg Argument of perigee u, deg Perigee distance p, km Time of perigee passage T, GMT 86.212637 20.742333 12.650441 90.421969 13.165104 90.439744 15:07:57 (28 Nov 1964) --41,535.874 1.1580740 28.133045 111.63727 245.65952 6565.7425 15:03:53.852 (28 Nov 1964) 2,022,402 3.1515691 142.00596 15.678090 186.01606 142.85406 89.557318 47.958569 1.2716438 270.00658 16:09:25 (5 Dec 1964) 228,218,340 30.785551 16,550.953 -- 5.1082557 289.33886 192.49794 30.007818 110.91233 0.05330023 269.96705 21:27:02 (23 July 1965) radial direction with decreasing speed; this reduced the geocentric angular rate of the spacecraft (in inertial coordinates) until the angular rate of the Earth's rotation exceeded that of the spacecraft. The phenomenon is illustrated (Fig. 4) on a map showing the Earth-track of the spacecraft reversing its direction from increasing to decreasing longitude. Also shown is the tracking sta- tion coverage and location of the various boost vehicle and spacecraft events. After several days of continuous tracking, it was esti- mated that, without a midcourse correction, the space- craft would pass the upper leading edge of Mars at a closest approach distance of 253,800 km. Closest approach would have occurred at 01:27:00 GMT on 17 July 1965. Comparison of these results with the desired Mars trail- ing edge pass indicated that the launch-vehicle's injection guidance system had performed within 3_ (three times the standard deviation) of the nominal values. 3. Mideourse Maneuver Phase To alter the trajectory so as to pass through a selected aiming region located at approximately 10,000 km from the center of Mars, a midcourse maneuver, utilizing a 16.70 m/sec velocity increment (87 m/sec capability), was required. In addition to altering the miss distance at Mars, this correction was selected to change the arrival time to 01:46:00 GMT on 15 July 1965, thus allowing the spacecraft's CC&S to activate various subsystems at the correct times near encounter. To properly align the thrust direction of the midcourse motor for the burn, a -39.16 deg pitch turn and 156.06 deg roll turn were required. The midcourse motor was ignited at 16:09:25 GMT on 5 December 1964, at which time the spacecraft was at a geocentric distance of 2,022,400 km and moving at an inertial speed of 3.139 km/sec relative to Earth. Analog data received at the Goldstone Tracking Station and relayed to the JPL Space Flight Operations Facility gave positive indication that the midcourse 6 JPL TECHNICAL REPORT NO. 32-881 d _C. '.;G'C . " (_) -_-_ ® iI -e o _j ® ® @ @ F @ @ I 7 JPL TECHNICAL REPORT NO. 32-881 maneuver and motor burn had been executed precisely. This was further verified by the observed doppler data, which were essentially the same as those predicted for the maneuver. 4. Cruise Phase Following the midcourse maneuver, the spacecraft reacquired the Sun and Canopus, thus returning to the cruise mode. At this time the spacecraft was moving primarily under the gravitational influence of the Sun in an ellipse with the Sun at the focus. During the early portion of the cruise phase, the spacecraft's heliocentric velocity was greater than the Earth's, causing the space- craft to lead the Earth around the Sun. This phenome- non is illustrated in Fig. 5, which contains a heliocentric plan view of the orbits of Earth and Mariner IV during the first 35 days of flight. Slowly, however, SUN EARTH \ ORBIT --'-_ 3C 2O 35 days ORBIT 3O 25 15 15 Mariner began to move out toward the orbit of Mars with decreasing heliocentric speed. On 28 February 1965 (61 days after launch), the Earth finally passed the space- craft in its orbital motion around the Sun. At this time, the spacecraft was at a distance of 18.9-million km be- hind the Earth and 320,000 km above the ecliptic plane. Throughout the rest of the flight, the Earth increased its lead in orbital rotation about the Sun (Fig. 6). Figures 7 to 13 present culwes of geocentric radius, geocentric speed, heliocentric radius, heliocentric speed, areocentric radius, Earth cone angle, and Earth clock angle as functions of flight time from launch to Mars encounter. Note in Fig. 12 that the minimum Earth cone angle (Earth-Spacecraft-Sun angle) was approximately 1 deg, rather than 0 deg, when the Earth passed the spacecraft in its orbital revolution around the Sun. (If the inclination of the heliocentric transfer orbital plane to the ecliptic plane had been 0 deg, the minimum Earth cone angle would have equaled 0 deg.) The heliocentric characteristics of the Mariner IV trajectory are shown in Table 3. Table 3. Heliocentric orbital elements of Mariner IV trajectory Pre-encounter Post-encounter Elliptical orbital element orbit orbit Semimaior axis a, km Eccentricity e Inclination to the ecliptic i, deg Longitude of ascending node _, deg Argument of perihelion c_, deg Perihelion distance p, km Time of perihelion passage T, GMT Period P, days 190,929,830 0.22750296 0.12569963 68.665534 352.56527 147,492,730 23:11:28 (23 Nov 1964) 526.64530 200,588,100 0.17322007 2.5437401 226.75545 200.64908 165,842,220 07:25:19 (16 Nov 1964) 567.11321 I0 SCALE: MILLIONS OF KM 5 0 10 20 0 days Fig. 5. Heliocentric plan view of Mariner IV trajectory during first 35 days of flight During the interplanetary phase of the flight, several orbital computations were made covering the period from the midcourse maneuver on 5 December 1964 to 10 July 1965 when the mass of Mars caused the first detectable perturbation in the Mariner IV trajectory. On the basis of these computations, it was determined that the closest approach to the surface of the planet would be 12,322 km occurring at 01:04:49.5 GMT, 15 July 1965. The areo- centric characteristics of the Mariner IV trajectory as predicted during the interplanetary portion of the flight are listed in Table 4. 8 JPL TECHNICAL REPORT NO. 32-881 ,,..C---" .--_" ?-" _"... \ \ 30 noys IL) 1 1 / I _( _'_ \ \ / t / -- -- Et )_ so .//// // _---- Iso leo\, \ ,, / 180 210 / 228 days / L_ ENCOUNTER JULY 14, 1965 Fig. 6. General relationship of Earth, Mars, and Mariner IV during cruise to encounter 280 24O E J< 200 "o C3 ,¢[ n- 120 n* I- z bJ 40 i j // J J J J _......__ i TIME FROM LAUNCH, days 225 250 Fig. 7. Geocentric radius vs time from launch Table 4. Areocentric orbital elements of Mariner IV trajectory Hyperbolic orbital element Se_lmaior axis a, km Eccentricity e Inclination to the ecliptic i, deg Longitude of ascending node _, deg Argument of periapsis _, deg Periapsis distance p, km Time of perlapsis passage T, GMT Pre*encounter prediction -- 22,046 0.65894 6O.458 188.O09 289.546 12,322 01:04:49.5 (15 July 1965) Actual Mars- encounter orbit --22,092 0.69753 58.186 187.499 289.321 13,201 01:00:58.1 (15 July 1965) 9 JPL TECHNICAL REPORT NO. 32-881 a 35 5O _25 E ;¢ c_20 I/J 15-- e_ F- Z E 280 240 a 200 I- z ta 160 120 ..._.__ j J J J J ENCOUNTER_-_, J J J 25 50 75 I00 125 150 175 200 225 250 TIME FROM LAUNCH, days Fig. 8. Geocentric speed vs time from launch J .__....... f" J J J J ENCOUNTER--,,_ _ ........----- 25 50 75 I00 125 150 175 200 225 TIME FROM LAUNCH, days Fig. 9. Heliocentric radius vs time from launch 25O 240 200 160 120 80 4O I00 ca 80 -o 6O 0 0 \ 40 2O 0 0 \ \ I ENCOUNTER 25 50 75 I00 125 150 175 200 225 250 TIME FROM LAUNCH, days Fig. 11. Areocentric radius vs time from launch \ \ 25 50 J ENCOUNTER 75 I00 125 150 175 200 225 250 TIME FROM LAUNCH, days Fig. 12. Earth cone angle vs time from launch _._NCOUNTER_ _ --------- _r---- Z5 50 75 100 125 150 175 200 225 250 TIME FROM LAUNCH, days Fig. 10. Heliocentric speed vs time from launch 5. Encounter Phase Mariner IV approached Mars along the trailing edge and from inside the planet's orbit (Fig. 14). Figure 15 illustrates the areocentric geometry of the flight past Mars. At about 43.5 min before closest approach, or at a distance of about 18,000 km from the planet's center, the Narrow-Angle Mars Gate sensed the planet, causing the tape recorder to start recording TV data 1 min 12 sec 360 300 240 180 120 tlJ 60 0 _,,_ ENCOUNTER_ 1 25 50 75 I00 125 150 175 200 225 250 TIME FROM LAUNCH, days Fig. 13. Earth clock angle vs time from launch later. At a distance of about 13,196 km from the planet's center, after a 25 rain 12 sec picture taking sequence, the scan platform moved permanently off the planet because of the angular movement of the spacecraft in its hyper- bolic orbit about Mars. At 02:19:11 GMT, the spacecraft 10 I JPL TECHNICAL REPORT NO. 32-881 ZONE OCCULTATION ZONE MARIMR TRAJECTORY RELATIVE TO MARS Fig. 14. Mariner IV and Mars orbits near encounter ON DENSITY, DEPTH Tv PIC ND MARINER FLIGHT PATH RELATIVE TO MARS TV PICTURES START, RKORDEO ON TAPE Fig. 15. Mariner IV encounter of Mars as seen from inside Mars orbit 1 1 JPL TECHNICAL REPORT NO. 32-881 passed behind Mars as seen from Earth (enter occulta- tion). The spacecraft emerged from occultation at 03:13:04 GMT after remaining behind the planet for 53 min 53 sec. Tracking data gathered and analyzed during the en- counter sequence indicated that the Mars encounter -- POST-ENCOUNTER TRAJECTORY \ \ t¢ I ON LAONC. /SEPT 8,1967 _ "_'_T 8' I_:ART H // SEPT 8, 967 /--PRE-ENCOUNTER , TRAJECTORY \ / Fig. 16. Heliocentric plan view of Mariner IV pre- and post-encounter orbits trajectory as predicted during the cruise portion of the flight was slightly in error. Orbit computations made as early as 1 hr before encounter revealed that the actual closest approach distance would be 13,201 km (instead of 12,322 km) occurring at 01:00:58 GMT. The principal causes of this discrepancy were an error in the opera- tional value of the astronomical unit (AU) and errors caused by small forces originating in the attitude control system. The areocentric characteristics of the Mariner IV trajectory as determincd from encounter data are shown in Table 4. 6. Post-Encounter Phase As it left the vicinity of Mars, the planet's gravitational pull altered the spacecraft's heliocentric orbit to such an extent that the perihelion distance changed from 147,493,000 to 165,842,000 km (Fig. 16). Because the spacecraft passed underneath the planet, a heliocentric- orbit-plane change of 2.67 deg occurred between the pre- and post-encounter orbits. To illustrate this effect, a plot of spacecraft celestial latitude as a function of calendar date is given in Fig. 17. By reference, again, to Fig. 16, it may be noted that the Mariner IV post- encounter orbit does not intersect the Earth's orbit. The minimum distance between these orbits, which represents the smallest possible closest approach distance be- tween the Earth and spacecraft, equals 16,244,000 km. Figure 18, which contains a plot of geocentric radius as a function of calendar date, reveals that the spacecraft will attain a close encounter with the Earth in 1967. "o U bJ / bJ U u i,i -3 ___ ENCOONTE.-- /I " /, I I I I I I I . I I I JAN MAR MAY JUL SEPT NOV dAN MAR FEB APR dUN AUG OCT DEC FEB 1965 \\ \ I / i / / \- / \\ / I J_ I I L I I I I I I I I MAY JUL SEPT NOV dAN MAR MAY JUL SEPT NOV APR dUN AUG OCT DEC FEB APR JUN AUG OCT DEC 1966 1967 CALENDAR DATE Fig. 17. Spacecraft celestial latitude vs calendar date 12 JPL TECHNICAL REPORT NO. 32-881 4OO 350 300 250 121 ne 200 I-- Z 150 ,o, ENCOUNTER- / ,oo j 50 J 0 l JAN. I MAR. FEB. // / I / f I I I NOVIMAY JUL. SEPT APR. JUN. AUG. OCT DEC. 1965 \ I t I _ _ JUL/ I IJAN MAR I. MAY t JUL/ SEPT NOV. JAN. MAR I MAY SEPT NOV. FEB. APR, JUN. AUG. OCT DEC. FEB APR JUN AUG, OCT. DEC, 1966 1967 CALENDAR DATE Fig. 18. Geocentric radius vs calendar date This pass, to occur on September 8, will yield a closest approach distance of about 47-million km. Curves of various heliocentric trajectory parameters for the post- encounter orbit are given as a function of calendar date in Figs. 19 to 22. The heliocentric characteristics of the Mariner IV post-encounter trajectory are shown in Table 3. B. Orbit Determination This section of the report presents a summary of the orbit results obtained from the Mariner Mars Mission. It will be divided into six categories. 1. Tracking Data Validity and Usage, containing a discussion of the high-frequency noise seen on the 500 E = 250 n_ I-- 150 z o_ _J _ I00 5O JAN. I MARl FEB. I I OUNTER MAY'juN. AU . OCT.DEC. 1965 JAN.I I IMAR. MAy -....___J J I I NOV.IdUE SEPI_ . J f_ FEB. APR. dUN. AUG. OCT. 1966 CALENDAR DATE , I 1 JAN/ MAR I. MAY i JUL SEPI: I NOV. DEC. FEB. APR. JUN. AUG. OCT. DEC. 1967 Fig. 19. Heliocentric radius vs calendar date 13 JPL TECHNICAL REPORT NO. 32-881 doppler tracking data, the actual data weights used for Mariner IV, and the rationale for using that weighting scheme 2. The Mariner IV Encounter Orbit Tracking [rom Encounter (E) - 5 Days to E + 5 Days, presenting a very accurate encounter orbit and giving both sta- tistical and non-statistical evidence for its validity (this orbit referred to as the true encounter orbit) 3. Evaluation of Combined Pre-Midcourse Orbit and Maneuver Errors, evaluating the combined error, based on the actual target point and the achieved target point (a preliminary effort to show how much of the total errors might have been due to orbit determination error) 4. Cruise Orbit Determination Errors, including dis- cussion and plots of errors (predicted target point 30oI 250 -- E ,.¢ . 200 a 15(3 _ iO0 o n_ 50-- 4 I I I JAN. MAR. FEB. APR I MAY i -- ENCOUNTER I i JUL. SEPT. NOV. / JUN. 1965 I I [ I I JAN. MAR. MAY JUL. SEPT AUG. OCT. DEC. / FEB. APR. dUN. AUG. 1966 CALENDAR DATE I I 1 I I I I NOV. JAN. MAR. MAY JUL. SEPt NOV. OCT. DEC. FEB. APR. JUN. AUG. OCT. DEC. 1967 Fig. 20. Areocentric radius vs calendar date 7O 60 50 z_40 (I 0 o 3O I p- ro 20 I0 ENCOUNTER-- _- I JAN. MAR MAY JUL SEPT FEB APR JUN AUG OCT 1965 \ \ , V NOV JAN MAR DEC FEB. J f i MAY JUL. f--\ • APR dUN. AUG. 1966 CALENDAR DATE i i I i MAY I JUL I. I ISEPT NOV JAN. MAR SEPT NOV OCT DEC FEB APR dUN. AUG OCT DEC 1967 Fig. 21. Earth cone angle vs calendar date 14 JPL TECHNICAL REPORT NO. 32-881 330 300 270 240 ._ 210 Z 180 o q o 150 -1- t- n- _ 120 9O 6O 30 I I I I 1 r¸ i I i i I I i I I I I i I _,ENCOUNTER I I t I dAN. MAR. MAY JUL SEPT. FEB. APR. JUN, AUG. OCT. 1965 I\II I I I I I t I I I I I I I 1 [ I I I I I I I I I i i i NOV, JAN. MAR. MAY JUL. SEPT NOV, JAN. MA_ MAY JUL. SEPt. NOV DEC FEB. APR. JUN AUG. OCT DEC FEB. APR. JUN AUG, OCT DEC. 1966 1967 CALENDAR DATE Rg. 22. Earth clock angle vs calendar date minus actual achieved target point) as a function of time from midcourse until 10 July 1965 (E - 5 days) 5. Near-Encounter Orbit Errors, presenting the real- time orbit errors from E - 10 hr to E + 5 hr 6. Conclusions, discussing the known error sources and giving a brief discussion of future investigations 1. Tracking Data: Validity and Usage Two-way DSIF doppler data comprised the only data type used in the orbits described here. Data were taken continuously during such critical phases of the mission as pre-midcourse maneuver and closest approach to Mars; data taken during other phases averaged 24 hr/5 day (these figures refer to 2-way doppler, only). Data acquisi- tion was from Goldstone, California; Johannesburg, South Africa; Woomera and Canberra, Australia; and Madr'.'d, Spain. The data time span considered here is from launch (28 November 1964) until 1 October 1965 (E + 2% mo). There is a reasonably good chance of ob- taining more tracking data when Mariner IV passes by the Earth in late 1967. These data were coherent without involving any use of microwave links to achieve coherency. No observable biases due to equipment were observed in the data, and the bias due to lack of knowl- edge of the refraction correction was minimized by allowing only data above 17 deg elevation in the orbit fits. High-frequency noise on the data was less than 0.02 cps for doppler counted over 60 sec. At the S-band frequency of 2295 Mc, this accuracy is equivalent to 0.12 cm/sec error in the range rate of the probe relative to the tracking station. It was possible to increase this resolution either by counting over larger counting intervals (accomplished by data processing of the original 60 sec sampled data) or by multiplying the doppler frequency to be counted by a factor (for Mariner IV a factor of 8) before the counting was performed. Thus, the data noise was generally < 0.003 cps S-band (0.02 cm/sec) as seen in the orbit residuals. The 0.003 cps 15 JPL TECHNICAL REPORT NO. 32-881 t represents the limiting accuracy of the existing single precision orbit determination program (ODP) to compute tracking data. The principal advantage to Mariner IV of such small data noise is in being able to see biases in the orbit residuals more clearly. Unfortunately, the physical model of the universe in the ODP is not exact, and this fact prohibits making the best use of the data. Since the ODP takes into account only those error sources formally implemented in a deterministic solution model and since the accuracy claimed by the ODP for its answers is di- rectly proportional to the accuracy the user claims for his data (assuming that no a priori information is used), the ODP for planetary cases would give grossly optimistic accuracy estimates if the input data accuracy estimate (data weight) were the same as the high-frequency noise on the data. This has been confirmed by actual experi- ence with Mariner II and Mariner IV. Furthermore, if the no a priori answers obtained from tracking data are combined with a priori constants which have reasonable associated statistics (such as the AU, station locations, and the solar pressure coefficients) as is actually done for most orbits, the a priori constant values will be out- weighed by the no a priori values. Thus, an error in the actual orbit itself can be induced by data weights that are too small. To avoid this as much as possible and to produce ODP statistics (tracking plus a priori data) that are reasonable, a semi-arbitrary range rate weight of cr_ = 3 cm/sec was chosen for 60-sec sampled data. Note that this is approximately 30 × larger than the noise seen on 60-sec counted S-band doppler and more than 150 × the noise on the 600-sec S-band data typically used in orbits. When 600-sec counted data are used (data points 600 sec apart) the _/N rule is used; this pro- duces a weight of _ = 8/k/10 _ 1 cm/sec. Essentially, this weighting scheme assumes that low-frequency biases due to program insufficiencies and physical model errors are the dominating error source and that data noise is so small as to be a negligible error source. It is obvious that using large data weights to compen- sate for orbit errors is not an optimum procedure. Un- fortunately, it is the only one currently available and, therefore, was used for both Mariner II and Mariner IV. The error sources for Mariner IV and the planned remedies for them are discussed under headings 4 and 5 of this section. In the absence of valid statistical descrip- tions of the orbit solutions, the labeling of solution parameters with statistical error estimates becomes a mat- ter of engineering judgement. In making those judge- ments, particular attention is paid to three factors: 1. The stability of a parameter being solved for when the conditions of the problem are varied (For ex- , , ample: (a) other important parameters can be frozen at various values and the effect on the parameter under test noted, or (b) solutions may be made using data from a single tracking station. By comparing several solutions of this type, a definite test can be made for the presence of bad tracking data.) The ability of the ODP to predict future measurable events such as Mars occultation times and future doppler observables The ability of the ODP to predict future events which are not directly observable but which are felt to be well determined after the conclusion of the mission (The near Mars orbit is believed to be such an observable.) 2. Mariner IV Post-Flight-Encounter Orbit Tracking from E - 5 Days to E + 5 Days This section presents a representative orbit computed after the flight on 30 October 1965. The epoch to which the initial position and velocity were referenced was on 10 July 1965 (E - 5 days). a. A priori I_ statistics input to the ODP. 1. Geocentric position and velocity of the probe, _po8 = 10,000 km spherical uncertainty _vel ---- 10 m/sec spherical uncertainty a priori values taken from cruise tracking 2. Mass ratio of Mars to the Sun Crm_lm s = 2% a priori value 0.32280422 × 10 -6 3. Solar pressure coefficient cr,+G = 5% _ 5% of the total solar pressure force of 2 dynes a priori value for i + G = 1.2067 4. Astronomical unit _AU = 2,000 km a priori value = 149,598,500 km 5. Station locations _po, = 50 m spherical a priori values from surveys, Ranger results and pre-Mariner IV encounter solutions b. Best available encounter orbit. When the solution was made, the geocentric position changed by slightly less than 1000 kin, the geocentric velocity by approxi- mately 1 m/sec, the solar pressure by 0.005%, and the station locations by less than 10 m. The solutions for 16 JPL TECHNICAL REPORT NO. 32-881 the AU and mass of Mars with their associated ODP 1_ statistics follow: AU = 149,597, 470 -*-240. and m_/,_ = (.322728 -*0.000015) X 10 -6 The encounter orbit parameters follow: E = Time of closest approach = TCA = 15 July 1965 01 h 0m 58._180 (GMT) B = 15_251.5 km *-15 B • Tc = 8,141.6 km ___20 B • Rc = 12,896.6 km -*20 Tracking data used in the fit was 600-see counted doppler data, except that 60-see counted doppler data were used from E - 4 hr to E + 4 hr. The important points are discussed in the following paragraphs. First, in a variety of other encounter orbits, the solu- tious for the encounter orbit (B, TCA, etc.) and the mass of Mars remained very stable. The range of values for the B °Tc and B. Rc components from the values quoted above was less than 5 km in all eases, the range of values about TCA was less than 2 see; and finally, the range of values for "x/_s from the quoted value was less than 0.01%. Second, for reasons probably connected with the sta- bility of the single precision computer program, a solution which included the AU always resulted in larger re- siduals between computed and observed data than a solu- tion that did not include the AU. However, the numerical value of the AU solution fell within a range of 200 km of the quoted value. It was shown that even if AU were frozen at a value of 149,598,500 km, which is 1000 km higher than the quoted solution, the target orbit and _/'s remained within the ranges given above. This would indicate, since AU is an ephemeris parameter, that ephemeris error does not affect the quoted solution sig- nificantly. Monte Carlo simulation studies, changing the look-up argument between Universal Time and Ephemeris Time by 20 see, showed similar results. Third, even for orbits fitting data from midcourse through E + 5 days, the range of values about B was less than 50 km, the range of _r/'_s less than 0.1%, and the range of AU less than 1000 km. This is surprising because, in most of these runs, no provision was made to solve for the Earth-Mars ephemeris errors and because the single precision ODP is not adequate for making such orbits. Fourth, the uncertainty in the radius of Mars obtained by astronomical measurements was approximately 50 km. The observed enter occultation time of the radio signal compared very favorably with that predicted by the encounter orbit and an average value of RM,,_ (3378 kin), thus indicating the orbit was valid to within 50 kin. Fifth, it is concluded from the preceding four points, that the encounter orbit has been well determined and _t_ t^,^..-_,, 1_ ,,_-._inH_ might reasonably be attached: as.to = B" Rc = 20 km a8 = 10 km O'Te A = 10 see amu/,.s = 0.02% of mu/ms aA_ = 500 km 3. Evaluation of Combined Pre-Mideourse Orbit and Maneuver Errors The aiming point for the rnidcourse maneuver made 5 December 1964 was: B = 12011 km B'Tc = 6007 km B. R_ = 10401 km TCA = 01 h 47 m The actual achieved target point from computation under the previous heading was: B = 15251 km B- Tc = 8142 km B-R_ = 12897 km TCA = 01 h 01 m Therefore, the errors were (actual minus desired) AB = +3240 km AB-T_ = +2135km AB" R_ = +2496 km ATCA =-46rnin The actual error was approximately three-fourths of the specified la requirement set by the Mariner 1V 17 JPL TECHNICAL REPORT NO. 32-881 Project. The la requirement on pre-midcourse orbit de- termination accuracy is given as 2250 km. This error is hard to assess since the estimated error is a strong func- tion of AU and solar pressure. However, preliminary results have shown that an error o£ -t-1,000 to 1,500 km inB'Tc and an error of +800 km in B'Rc to be the most likely pre-midcourse orbit error. Much work remains if more definite numbers are required. 4. Cruise Orbit Determination Errors Orbits were performed at an average of one per two weeks during the cruise phase (midcourse to E - 5 days) of the mission. During the first few months of the flight, the data were not strong enough to solve for AU so that, essentially, the AU was frozen at the JPL Venus radar bounce value of 149,598,500 (±500 km la). Later it was possible to loosen the a priori on AU somewhat, allowing the data to participate more fully in determining AU. Since the solution for AU tended to change somewhat from week to week, there was a corresponding jumpiness in the target parameters, which was intensified when, approximately one month before encounter, the mass of Mars was added to the solution set. Thus, there was no set of runs done in real time during cruise that used the same a priori assumptions. The plots (Figs. 28 to 25) of real-time orbit errors show trends consistent with the above procedures in that the orbits are stable for the first few months and then tend to jump around late in the mission. Two other curves are also shown in each o£ these figures. These curves repre- sent systematic sets of runs done after the end of the mission in which all a priori conditions were held con- -200 I-- -400 Z -600 -800 v -I000 Ld -1200 -I 400 l / / i / I JAN i 1965 ® FOR AU = 149597500 km \ • FOR AU= 149598500km - •". FOR REAL-TIME RUNS I I .L FEB I MAR I APR I MAY DATA FROM MAR/NERJZ_-, da -[ - _ CLOSESTAPPROACH TO MAR _ i--HI, I dUN JUL Fig. 23. Mariner IV errors in predicting B'Tc 200 -- F- E 0 __2oo ( ---------¢- _.oo'e -600 -- " -800 SCH / | ® FOR AU= 149597500 km '_"-'0°°1_ F • FOR AU= 149598500 km _,200_ AL-TIMERUNS I dAN I FEB I MAR I APR I MAY JUNE I I0 JULY 1965 DATA FROM MARINERI_', de Fig. 24. Mariner IV errors in predicting B" Rc 5OO 480 I--- co 360 a 240---- u..J I-- (..) (:3 LLI n_ i20 --'-- 0 -120 I JAN I FEB 1965 I I I ® FOR AU = 149597500 km • FOR AU = 149598500 km -- ,', FOR REAL-TIME RUNS k I I t A-_ CLOSEST APPROACH I TO MARS- [ MAR APR MAY I JUNE I JULY I AUG DATA FROM MAR/NER.ZE, doy Fig. 25. Mariner IV errors in predicting time of closest approach stant and data were added in 1-mo blocks. For one set of runs, AU was frozen at the JPL Venus bounce AU value of 149,598,500, and the other set AU was frozen at the value obtained from the Mariner IV encounter fit (149,597,500). All available features of the program were used in these latest runs, including improved timing equations, an improved estimate of the solar pressure from telemetry, and the use of solar pressure forces in two components normal to the Sun-probe line. These K/r_' type forces were needed to approximate the forces induced by moving solar vanes on the ends of the solar panels. Since the telemetry measurements of these vane positions gave better knowledge of the normal solar forces than was possible by solving for these forces with track- ing data, the corresponding constants were included in the trajectory but not the parameter solution list. 18 JPL TECHNICAL REPORT NO. 32-881 When AU = 149,597,500 km is used, the remarkable reduction of target prediction error enhances the validity of the otherwise suspect encounter AU solution. Other recent results from Venus radar bounce observations give a result of AU = 149,598,000 km. Therefore, since _B/_AU = 0.6, it is likely that the real-time AU value of 149,598,500 km produced a target error of from 300 to 600 km. As for the real-time runs, it is seen that there are errors of from 500 to 1000 km throughout cruise, even though the consistency of the orbits is high for certain portions of the flight. 5. Near-Encounter Orbit Errors (Epoch 0h 10 July 1965; Tracking Data from Oh 10 July) This section presents the errors in the orbits performed starting at closest approach, minus 10 hr. Note that the errors at that time are roughly those existing at the end of the cruise plots (Figs. 23 to 25). A priori sigmas were possibly somewhat tighter on position and velocity than would seem wise in the light of post-flight analysis. However, during the mission, there had been no evidence of an error of the size that actually existed. During the last 5 hr before encounter, two computers were opera- ting. One computer was using "_'pos = 500 km and "b', = 0.1 m/see, while the other used %'po, = 200 km and "_', = 0.1 m/see. Later, at approximately E -- 1 hr, both computers used the looser a priori. The errors in several representative orbits are shown in Fig. 53. Con- siderably more work will be done in making systematic runs similar to the two cruise sets of runs described in the preceding section. Operationally, the orbit errors during the period from E - 10 hr through E + 5 hr did not result in any degradation of the scientific results from Mariner IV including the television pictures and the Mars occultation data. 6. Conclusions a. Adherence to specification. Operationally, the orbit determination error met the error specifications set by the Mariner IV project and did not degrade the per- formance of the mission significantly. b. Deviation from predicted errors. The orbit errors during the cruise and near-encounter portion of the mission were 3- or 4-times larger than had been expected. 40O ZOO E o rr h- _,0o' n_ 0_ _-_o 0 n- 5 -8o6 -I00( -I0 i _B'Tc _ ! I I -8 -6 -4 -2 i ! i / i/ i ! i r E I ' i I I J E 2 4 6 TIME FROM ENCOUNTER, hr Fig. 26. Errors in predicting encounter quantities vs time, Mariner IV Some of the possible causes, or multiple causes, are the following: 1. ODP computations are done mainly in single-pre- cision (8-digit) arithmetic. This cattses some insta- bility in the cruise orbits but is not a major error source. 2. Solar pressure forces caused by solar vane positions were not implemented in ODP until after encounter. These forces cause a 200-km error. 3. Telemetry data indicated random forces of 0.5% of the solar-pressure force (or larger) due to valve leakage from the attitude control system. The leak rate was found to vary occasionally when a valve was actuated. From the limit cycle telemetry, it is possible to verify frequent changes in the leak rate of 0.5% of the solar-pressure force. The solar- pressure force averaged 4 dynes on the 260 kg Mariner IV spacecraft. The time-averaged transla- tional force due to the total leak rate is greater than this but very probably no greater than 3% of the solar-pressure force. Pre-flight measurements of leak rate on the Mariner IV valves typically showed leak rates in the range of 0.5 to 2_ of the solar pressure force. One perc_ent of the solar pressure force maps into a B-plane error of 200 km. 4. Planetary ephemeris error is a possible error source, although the ephemerides are believed to be ac- curate to 400 km (Sa). 19 JPL TECHNICAL REPORT NO. 32-881 c. Difficulty in fitting data. In the light of the above mentioned error sources, it is understandable that the cruise data were difficult to fit and that the residuals should show long-term trends with maximum amplitude of 0.05 cps S-band (0.3 cm/sec range rate). Taking this disturbance into account, it is also plausible that the stability of the cruise orbit results might deteriorate when two highly correlated effects such as solar pressure and the AU are solved for simultaneously. d. Program Jot complete analysis. A definitive analysis of the Mariner IV orbit errors cannot be completed until proper computer programs are available to analyze the data. Such analysis will require the new double precision orbit data program (DPODP), which is scheduled for completion in early 1967. In the meantime the investiga- tion will be carried out with improved versions of the current ODP. There is considerable doubt as to the extent to which the attitude control forces can ever be repre- sented in the solution, even with the DPODP, because of their non-deterministic character. e. Solutions for mass oJ Mars and AU. The mass of Mars and AU solutions are reasonably consistent with earlier experimenters. In Figs. 27 and 28, bar graph plots show the various values (by the dark lines) and the probable errors (by the length of the bar away from the value line). All solutions except that for Mariner IV use Venus radar bounce data taken near the Earth-Venus conjunctions. No value is given for the Mariner II 1962 Venus flight; the definitive Mariner II AU solution will probably be published in the next year. For Mariner II, the encounter AU solution was a factor of 25 weaker than the Mariner IV encounter solution; thus, to obtain a stronger Mariner II AU solution, the cruise and en- counter data must be combined in one orbit. The delay in achieving a high-accuracy Mariner II AU solution has been caused by difficulties in making the combined solu- tions. The Mariner IV combined solutions will also be released within one year. The JPL Venus bounce value, to to to Zrr _<_ 5,120,000 5,100,000 3,080,000 3,060,000 3,085,000 3,110,000 3,079,000 3,098,600 :1:5,000 +7,700 :1:5,702 :1:600 de SITTER RABE UREY JPL 1938 1949 1952 1965 WEIGHTED EROS DEIMOS MAR//VER MEAN Fig. 27. Mass of Mars solutions which is marked "provisional" (not yet published), is quoted with the permission of Dr. D. Mtuhleman of JPL. This solution is considered to be more accurate than the 1961 JPL value since it uses a double-precision computer program, solves for the Venus ephemeris as well as the AU, and has data from three coniunctions. The 1961 solu- tion was in single precision and did not solve for ephem- eris corrections. All values are for a value of the speed of light of c = 299792.5 km/sec. The speed of light is considered to be a defined constant in orbit and plane- tary bounce work, so that lengths are really determined in light seconds and are then translated to meters for more convenient use. 2O L__ JPL TECHNICAL REPORT NO. 32-881 J JPL MARINER .Z]Z"PROVISIONAL -----_ JPL 1961, 1962, 1964 DOPPLER ONLY PRELIMINARY I USSR 1964CONJUNCTION USSR 1961 ICONJUNCTION JPL 1962 CONJUNCTION JPL PUBLISHED 1961 CONJUNCTION 7000 I I t I 7500 8000 8500 9000 (ALL NUMBERS ARE VARIATIONS OF 149, 590,000 ) AU SOLUTIONS, km Fig. 28. AU solutions from Venus bounce and Mariner IV data i 9500 I0000 21 JPL TECHNICAL REPORTNO. 32-881 III. SPACE FLIGHT OPERATIONS A. Data Recovery and Processing The data recovered from Mariner IV was, basically, of two types-tracking and telemetry-which may be related, primarily, to two phases: (1) the high-activity times of launch, midcourse maneuver, and encounter; and (2) the cruise phase. Other periods of high interest will, also, be discussed. It was expected that hardware and software problems would be encountered in this thorough operation of the SFOF Data Processing System used throughout the Mariner IV mission; however, the majority of these were resoh,ed. Certain system anomalies were uncovered and changes are being incorporated into the design. 1. Science and Engineering Data From liftoff to antenna changeover, the Mariner space- craft transmitted approximately 3 X 10 s bits of data back to the earth. Of these, 107 bits were television data, 97 X 10 '_bits were engineering data, and 193 X 10 '; bits were science data. Approximately 98_ of the data transmitted to the SFOF in real time was processed and logged by the IBM 7040 computers. The IBM 7094 post-processing recovered approximately 90_ of the available data. The 2_ lost in real-time processing was due to communication out- ages and computer outages. The 10% lost by the 7094 post processing was caused, primarily, by the inadequate assignment of time with data and by the inability of the edit program to identify preambles. The Master Data Library (MDL) effort recovered from 99.5 to 10f_'_ of the data recorded at the DSIF stations. Of the total data transmitted from the spacecraft, about 85_ was actually presented as the MDL; this percentage includes data missed during the two periods that Station 51 was net tracking the spacecraft, as well as all that lost as a result of individual station outages. 2. Tracking Data The number of tracking points taken during the Mariner IV mission may be estimated from the following tracking coverage summary: 1. Elapsed time was 807 days between 29 November 1964 and 1 October 1965. 2. Total tracking coverage was approximately 9500 hr, averaging 10 hr/day per station for three stations. 3. The tracking pattern from launch to encounter com- prised 1-sec, 10-sec, and 60-sec data. (a) A total of 20 min of data were acquired during the following periods of each pass: from horizon (rise) to horizon + 20 min; at maximum elevation + 10 rain; and at horizon - 20 min to horizon (set). (b) The 60-sec data were taken at other times during each pass. 4. Two passes per week of horizon-to-horizon 2-way data were taken each week per station. The 60-sec points totaled approximately 550,000; the 10-sec points approximately 320,000; and the 1-see points approximately 3,200,000. The 60- and 10-sec points were sent in real time to the SFOF. It is estimated that close to 1,000,000 points were transmitted to the SFOF in real time, of which approximately 300,000 were precision two-way points. The 1-sec points were shipped in non- real time for Master Data Library post-processing. 3. Coverage a. High Activity Phases. During the high activity phases of the mission, redundant systems were utilized wherever possible. At the SFOF all operational com- puters were up and on line. Both the 7040 and 7094 computer strings were operated in computer Mode II. _ During the critical phases, the 7094 data processors were divided between the two computer strings. One 7094 was used primarily for flight path programs; the second was used for engineering and science programs. All pro- grams were loaded into both strings so that an immediate backup was available for any given program. Continuous orbit updating was performed from launch to launch + 20-hr. After that time, orbit updates were performed approximately every 6 hr up to midcourse maneuver; then, tracking data monitoring was done on a nearly continuous basis. Because of the unreliable performance of the Data Processing System (DPS) in handling tracking data, it was decided to back up the DPS by punching cards with tracking data information. These cards could be read 'Computer Mode II consists of a 7044 (input/output processor) linked to a 7094 (detailed analysis machine) by a data storage disk. 22 directly into the 7094 Tracking Data Processor (TDP), and merged with data coming into the 7094 via the 7040. This backup procedure retrieved much valuable data that, otherwise, would have been lost in the early hours of the mission from various malfunctions in the DPS; during the first 30 days of the mission, only about 10 or 15% of the data were processed through normal channels (7288 data channel multiplexer with the 7040 and 7094 computers). When the card information was added, this figure increased to about 85 to 90% of the usable track- ing data. Corrections made to the mission-independent editor in the 7094 increased the data recovery to about 85 to 90% through the DPS; it remained at this figure, or better, throughout the remainder of the mission. b. Cruise Phase. During the cruise phase, only the 7040 computer was used in real time. The real-time computer output consisted of: (1) Quick-look formats for science and engineering, (2) high-speed teleprinter formats of selected engineering measurements, and (3) 30 × 30-in. plots of selected science and engineering measurements. During this time, the 7040 computer also generated a log tape consisting of all the incoming raw data. No real-time processing of tracking data was done in this mode. The 7094 user programs were run on a production type operation. The 7040 log tape was removed at 6:00 a.m. each day and edited by the 7094. Science and engineer- ing programs were then run on the edited data. The computer printout was duplicated by document control and then delivered to the user areas. User program out- put for a 24-hr period was delivered to the user areas by noon the day following its acquisition. During the cruise phase, tracking data monitor runs were made to validate the tracking data quality approxi- mately every two or three days. At these times, the track- ing data file was updated with the new data. Orbit updates were performed every two weeks during the cruise phase. 4. Ground Telemetry Subsystem Pedormance The Ground Telemetry Subsystem for Mariner con- sisted of the demodulator, the decommutator, and the TrY encoder. The system performed satisfactorily throughout the Mariner mission. The output of the de- modulator was occasionally data bar (compliment of data), however, the problem was easily corrected by re- synchronizing the demodulator. The data bar condition was most prevalent during station acquisition. JPL TECHNICAL REPORT NO. 32-881 The TTY encoders performed well throughout the mis- sion. However, two anomalies were observed at the SFOF: (1) When the station paper tapes were processed by the Telemetry Processing Station (TPS) at the SFOF, excessive skew was evident. The skew was generated by the punches in the TrY encoders at the stations. Me- chanical readers were not affected by this skew; how- ever, TPS used an optical reader for increased speed of reading. (2) During TV picture playback, the TTY en- , • ___._.. 11.. • W¢o cartage re_ha_n_ intocoder WOUlU occu lo.any h_c_ " _ _ the data stream instead of one. This caused a problem for the data processing system at the SFOF and it was necessary to modify the mission-independent editor to accommodate the situation. No problems affecting data recovery were encountered with the decommutator during the mission. 5. Picture Data During the encounter and the beginning of picture playback, full data processing capabilities were used. Both computer strings were exercised, one for flight path and the other for science and engineering. After the playback of picture 1, the computer coverage was re- duced to two 7040s making redundant recording tapes, and 7094 video processing was performed each morning at 6:00 a.m. The 7094 processing of the picture data produced two outputs. One was a computer printout, and the other was a magnetic tape. The 7094 computer printout consisted of a number of formats that displayed picture data, as well as summary data, concerning the picture. The picture data were dis- played in a 200 × 200 decimal element matrix and, also, in a printout of the binary serial bit stream. The magnetic tape was written in a manner which was compatible with the link film recorder. The above processing constituted the SFO commit- ment to the Project; however, an extensive enhancement effort was conducted by Space Science personnel. 6. Occultation Phase As Mars encounter approached, the orbit determina- tion effort was increased. Daily orbits were completed during the two weeks prior to encounter. Continuous orbit runs were made on one string, and periodic orbit runs were made on the backup string from closest 23 JPL TECHNICAL REPORT NO. 32-881 approach - 12 hr to exit occultation. During the occul- tation period (enter occultation - 30-min to exit occultation), residual data generated by differencing the incoming data and predicted data were transmitted to the SFOF from the Goldstone Echo Site and plotted in near real-time on the FPAC 30 X 30 plotter. Unexpect- edly large residuals, from difficulties with the orbital calculations at encounter, caused part of the plotting effort to be of limited value; however, it was possible to verify the existence of the Martian atmosphere and make some preliminary estimates of its character in near real-time from observation of the plot. The delay in lock- ing up the closed loop at Goldstone at exit occultation made it impossible to plot residuals at that time. 7. Mariner Master Data Library Since August 1964, an active program was underway to design, develop, and implement a system to produce a master library of all data received and recorded during the Mariner Mars 1964 mission. This Master Data Library (MDL) would provide the best source of telemetry and tracking data received and recorded by the DSIF during the Mariner mission, as well as discrete DSIF instrumen- tation performance parameters. The primary purpose of the MDL was to produce a history of the Mariner mis- sion from which post-flight analysis could be accom- plished on the spacecraft's subsystems, the scientific payload, the spacecraft's trajectory, and the performance of the DSIF. To afford a more flexible and expedient access to data comprising the MDL, the system was categorized into the three data tables-telemetry data, tracking data, and comment data. a. Telemetry Data. The MDL telemetry data table comprised two types of digital-recorded magnetic tapes generated by IBM 7094 computer programs: Station Master Merge Tapes and Composite Master Merge Tapes. Station Master Merge Tapes. The Station Master Merge Tapes were generated by the Merge Program for each Deep Space Station (DSS) and contained the best telemetry data and ground instrumentation performance parameters recorded dnring each station's pass. The best telemetry data available was derived from the three data sources processed, i.e., the demodulator input, the de- modulator output, and the teletype encoder output. The selection of the data source was made by the Merge Pro- gram based on the data quality and continuity of each of the data streams within a recorded source. Composite Master Merge Tapes. The Composite Mas- ter Merge Tapes, which were also generated by the Merge Program, contained a continuous and sequential stream of the best telemetry data derived from the com- posite of DSIF stations receiving and recording telemetry data (Station Master Merge Tapes) during each day of the Mariner IV mission. The overlap of telemetry data between DSIF stations was eliminated by data source selection based upon data quality and continuity of each recorded source. The ground instrumentation perfor- mance parameters were not contained on these tapes. b. Tracking Data. The MDL tracking data table was developed from both real time and non-real time data. Real-time data (data sampled at low rates, only) were received from the DSIF stations via TTY communica- tions into the SFOF and were processed through the DPS in real time. Non-real-time data (data sampled at both low and high rates) were received from the DSIF sta- tions and were, also, processed through the 7094 portion of the DPS. Both real-time and non-real-time tracking data were maintained on digital recorded magnetic tapes in formats compatible with the TDP Program and Orbit Data Generator (ODG) Program, which were in them- selves user programs. The raw tracking data, as recorded on TTY paper tapes, were maintained as part of the MDL. c. Comment Data. To explain peculiarities and/or anomalies that occurred in both the telemetry and track- ing data tables, the supporting flight operations logs and MDL processing logs were maintained in the comment data table which was recorded on microfilm. The flight operations logs consisted of the SFO log and the DSIF Operations log from each DSS. The history of the MDL processing was contained in the MDL data processing logs and the actual telemetry and tracking data which were printed in tabular outputs. Each type of operations log, MDL data processing log, and tabulated output was maintained on separate positive-microfilm records pro- viding convenient reproduction and access by users. B. Space Flight Operations Performance During the launch phase of Mariner IV, the DPS, which was configured for two Mode II systems, encoun- tered no problems. Approximately one-half of the period frcm launch to maneuver was devoted to Mode II pro- cessing and the other half to Mode IIF processing. The only significant failure during this period was that the 7040 occasionally had problems processing high-speed time words, which caused a loss of data synchronization. :In computer Mode III, the 7044 operates alone. 24 JPL TECHNICAL REPORT NO. 32-881 MARINER MISSION SUPPORT AREA _ NASA Pt_uJECT bi_AGER 5439 ACE 2 SOFT MAN ACE I 5843 5842 5839 5841 5336 5837 MISSION MISSION CONTROL CONTROL CONTROL ROOM ROOM ROOM ]1 I I Fig. 29. Operations area 5 84 5915 _ _ _5913 I 912 , ti 15 °5 T 5909 5921 I 5870 FLIGHT PATH ANALYSIS AREA I _ 5918 Rg. 30. Spacecraft performance analysis area Z 1_ Two Mode III computer systems were maintained throughout the maneuver operation with no significant problems or failures. Figures 29 through 32 show the configuration of the analysis areas provided in support of Mariner IV from launch through maneuver. Figures 33 and 34 show the areas provided for support of Mariner throughout flae erie pha_e aud unHj mis- sion termination on 1 October 1965. Four Operations Controllers maintained 24-hr/day alarm monitoring of the Mariner IV spacecraft during the cruise portion of the mission. Some of the duties per- formed by the Operations Controllers were the following: 1. Alarm monitor the incoming real-time data. 2. Notify cognizant personnel should a spacecraft alarm condition exist. 3. Monitor data condition. 4. When incoming data is not good, determine the cause, communication line, or computer, and take corrective action. [ I SCIENCE CONFERENCE ROOM 5895 I II IsP,CEII 3_8:_0 PLOTTER _ 5891 5894 I 5914 [ 5898 L S ,3TV Fig. 31. Space Science Analysis area, first floor 25 JPL TECHNICAL REPORT NO. 32-881 5. See that data distribution by the Technical Assis- tants is accomplished daily. 6. Update status board daily. 7. Update status phone daily. I [ L_ 5952 SPACECRAFT PERFORMANCE ANALYSIS AREA TRAJI 59291 __._._.._ 59291 m 5934 59511 58151 FLIGHT CHIEF 5925 r---] FLIGHT 5926 1 5938 5925 5901 I ORBIT 5854 TRACKING 5955 TRACKING DATA ! ANALYSIS I 5956 Fig. 32. Flight Path Analysis area, first floor An Engineering Planner was assigned to support the project throughout the entire mission. The functions per- formed in support of the Mariner IV mission were: 1. Produce Mariner Mars 1964 Technical Bulletins. 2. Schedule facilities and computers. 3. Maintain flight history chart. 4. Provide inputs to documentation. 5. Provide prelaunch test scheduling and distribution of test notices. 6. Maintain master file of Mariner documentation. 7. Write Mariner failure reports. 8. Publish daily status meeting minutes. CONFERENCE ROOM 5965 I 599 4 59901 6165 5959 L5655 I I 5245 I SCIENCE 'till I I I I I SCIENCE BUSS CHIEF 5951 SPAC DI RECTOR 5831 BUSS II 5967 RADIO _ TELE I 59,9III 58,B COMMAND _-_ TELE n C_R I_ _ 1 '_ )NSOLE 5_ 69 I 5962 965:/o GUIDANCE CONTROL 5875 5903 I_TEM PE;6AT URE hl ' 1 Fig. 33. Mariner Mission Support area 26 5881 I MARtNER MISSION SUPPORT AREA 5R7cl I _J 5668 \--TRACKING DATA ANALYSIS I I t I 5871 TRACK 2 DSIF NET CONTROL 5908 1 E:= LOBBY --.11,,. 5882 ] I , Fig. 34. Track area 9. Act as standby Operations Controller. 10. Assist in maintaining SFO Flight Log. During the cruise phase, the data processing system underwent a series of changes designed to increase the capacity and flexibility of the system. The basic moves made in this period were the installation of a 7044 in place of the 7040 in the X string, the modification of the 7094 in the Y string to be a Model II, and the move of the Data Processing Control Area to the south end of the Computer Area of the SFOF. The completion of these moves took place on schedule and with little or no complication in the performance of the work. The installation of this equipment was followed by a period of trial during which many problems were found in the remainder of the system. l. Investigation and Correction of Problems Although some of these problems were still being in- vestigated and solved, the majority of them were cor- rected prior to encounter. The week of testing prior to Mariner encounter revealed many of the problems and resulted in a concentrated effort to get them cleared up in time for the encounter; therefore, the week of encoun- ter produced an exceptionally clean period of operation. JPL TECHNICAL REPORT NO. 32-881 There were a few minor problems but nothing of opera- tional importance. At the end of the period, however, one of the old installation problems recurred. A summary of the problems encountered during the period of this re- port follows. a. Memory errors. A memory driver failure was iso- lated and repaired in one day. Memory parity errors appeared frequently on both 7044s. These errors were apparently cured after a team from the IBM factory in New York spent a week prior to Mariner encounter work- ing on both 7044s. During this time, they ran digital and analog diagnostics and did considerable adjustment of the memory circuits which affected the more critical timing of the high-speed memories. At the end of this period, there seemed to be no more memory parity errors. The computer system performed very well through en- counter, but on 15 July there was another memory parity error on the 7044X. This was followed during the week of 21-28 July by a series of memory parity errors on the 7044Y. These parity errors were found to be due to two bad electronic circuit cards in the memory parity register. Card replacements corrected the major portion, if not all, of the memory parity error problems. Most direct data channel failures were corrected after a short period of observation and trouble shooting. One remaining problem was found during the week of 5 July in an encounter testing period and was corrected after about 5 hr of down-time during the test. b. Wiring problem. During the investigation of the memory parity problem, it was discovered that the com- puters had not been wired properly to the power lines. The main source of the problem was the substitution of the neutral wire of the three-phase power system for a green ground wire. This caused the computer equipment frame ground to be noisy, reflecting the voltage drops caused by the current-carrying neutral wire. The problem was corrected. c. Connector wear. A continuing problem had been the inability of the connectors on the 7242 M-box to with- stand the constant reconfiguration of the input/output (I/O) devices. This constant reconnection caused the pin that retains the shoulders in the connector block to be- come worn, causing pushed pins and loss of signal con- tinuity to the device. d. Other problems. Although there was the usual amount of routine problems, in addition to those outlined above, they were of little individual significance. They 27 JPL TECHNICAL REPORT NO. 32-881 consisted, principally, of such items as data processing card reader failures which were easily diagnosed, or such others as tape drive failures; these were corrected easily and in a routine manner. A failure of the under-floor insulation of an M-box ground wire caused the M-box to become nonoperational and prevent function of the input/output of the 7044s. This problem, massive in its symptom but subtle in its analysis, was corrected in about 2 hr. 2. Three-Shift Coverage, 6 July to 6 August 1965 Support of the computer subsystem starting the week of July 6 included third shift coverage as well as the usual two shifts. This coverage continued until 6 August. A third SC-3070 was installed in the Mariner control center and a 30 × 30 plotter in the Operations Area for encounter. No severe problems were met in the I/O equipment in the User Areas. The I/O operators were scheduled for continuous 12-hr shifts from 6-9 July (the pre-encounter test week) and from 12-16 July. The I/O maintenance personnel were on continuous 12-hr shifts from 6-10 July and 13-16 July. This coverage, and the spotting of stand- by equipment in the Operations Area, enabled equipment changes on almost a minute-by-minute basis, which in several instances allowed User Areas to remain on-line w_th virtually no interruption of service. On 1 July 1965, a short-duration power failure in the Edison commercial power line caused an automatic shut- down of all computers. After a short period of waiting to ensure that the power transient would not be repeated, the computers were brought back up, diagnostics run, and the operating programs reloaded. This procedure took about 1 hr and 20 min. All communication subsystems within the SFOF sup- ported the Mariner Mission. 3. Support Services Rendered The SFOF Operations Support System and the Dis- play System provided the following assistance to the Mariner IV mission: 1. Maintenance of electrical and air-conditioning systems 2. ttandling of space flight data for distribution and storage in SFOF document control 3. Logging 4. Scheduling 5. Mariner Mission Support Area technical assistance 6. Information display 7. Access control 8. Special services 9. Documentation Power supply and air-conditioning were maintained. In order to provide for any emergency situation due to loss of power, a standard operating procedure (SOP) was written and authorized. Instructions for handling power failure in order to minimize a possible loss of data were given in this SOP to all personnel participating in mis- sion endeavors. a. SFOF Document Control. The SFOF Document Control group provided the following support for the Mariner IV Mission: 1. Data Accumulation Magnetic tapes Station tapes Master data library (MDL) tapes Time sequence tape of best data Station overlap tape of best data Teletype tapes 2. Microfilm Rolls Processed 16 mm All computer program printouts 35 mm Communication center page printout Oscillograms 3070 printout SFOF Operations Logs SC 4020 3. Binder formatted printouts of computer programs for analysis groups 4. SFOF Operations Log 5. Station logs 6. Sequence of events for both tests and operations In addition, Document Control handled data distribu- tion to all the areas of the operation, which were further distributed by the Technical Assistants. Document Con- trol provided service for recalling data for any portion of the Mariner flight upon request. 28 Thesystemusedby theSFOFDocumentControlin supportof the missionperformedsmoothlyand effec- tively. Thetypingof SFOF Operations Logs for Mariner was performed throughout the mission. These logs, which provided a continuous record of SFOF operations, were microfilmed for inclusion in the Master Data Library. Technic*_] A_sistants were provided to Operations throughout the mission. Among many other things, they provided computer printout distribution as shown in Table 5. MSA bulk data distribution JPL TECHNICAL REPORT NO. 32-881 The real-time AGC display on the Sanborn Recorder in the MMSA displayed data from all prime DSIF sta- tions supporting Mariner. b. Access Control. The Support Group was supplied with Mariner-dependent access lists, and communication, DSIF, and data processing lists of personnel to be given access to closed areas of the SFOF. The lists were computer-generated. Badges were issued to 650 people. Control was maintained during tes_g and acbJa! operations. Security was enforced at the access desk and by six guards. Table 5. Computer printout distribution EAPM, PAP Power analysis 1 CCAP Central computer & sequencer 1 ACAM, ACAP Attitude control analysis 1 CPPM Communications predictions 1 Plots " AGCM Automatic gain control Plots ATRM Attitude reference 1 Plots ENGR Engineering data display SSDM System suppressed data 1" EPLM Plots (separated) ",P SIPM Star data, magnetometer antenna dafa 1 Plots VIPM Variable parameter input aReproduced copy (81/2 X 11) within 2 to 4 hr. bUnseparoted. i 1 1 1 1" 1" 1 2 I" I l a I" I a I" 29 JPL TECHNICAL REPORTNO. 32-881 The SFOF Operations Support Group arranged for special services for the Mariner mission including the following items: 1. The handling of a special sequence of events (Appendix D) for encounter. 2. A form for DSN communications coverage. 3. A special telephone directory (300 copies) of Mariner Mars 1964 mission-dependent personnel. This directory listed personnel, their titles, SFOF extensions, and Laboratory extensions. The SFOF Documentation Group prepared and pub- lished the documents listed in Table 6 in support of Mariner IV. Table 6. SFOF Mariner documents a. Standard Operating Procedures a. (Cant'd) SFOF/SOP Title Date No. SFOF/SOP Title No. M20-300A M20-302 M20-306 M31-105 M31-320 M40-100 M45-100 M45-101 M45-102 M45-103 M45-110 M45-111 M45-120 M45-130 M45-131 M45-132 M45-133 M45-134 M45-135 M45-136 M45-137 Date DSIF Predicts Transmission 10/30/64 Punching Tracking Prediction Tape 1/21/65 Mariner Tracking Predicts for Non-DSN 4/2/65 Tracking Stations Special Instructions for Teletype Patching 9/24/64 Mariner Midcourse Maneuver Availability of SFOF Conference Nets Outside 10/16/64 SFOF Operations Ai'eas Computer Operations Power On/Off 4/6/65 Computer Operations: 7040 Computer 10/9/64 Operations Computer Operations: 7094 Computer 10/9/64 Operations Computer Operations: M-Box Setup and Channel 10/9/64 Override Switch Setup Computer Operations: Responsibilities of 10/9/64 Personnel Computer Initialization: 7040 Initialization 10/9/64 Deck Parameters Computer Initialization: 7094 Initialization 10/9/64 Deck Parameters Editor Operating Procedures 10/9/64 Operation of the DPCC: General Responsibilities 10/9/64 Operation of the DPCC: Hardware Preparation 10/9/64 Operation of the DPCC: Software Preparation 10/9/64 Operation of the DPCC: Data Stream Status 10/9/64 Operation of the DPCC: System Start Up 10/9/64 Operation of the DPCC: Data Stream Selection 10/9/64 and Data Identification Operation of the DPCC: Special DPCC Functions 10/9/64 Operation of the DPCC: System Performance 10/9/64 Monitoring M45-138 M45-140 M45-200 M45-201 M45-202 M45-203 M45-204 M45-300 M45-350 M70-115A M72-105 Switching Trap Procedure Computer Operations: 1401 SFOF Output Program TPS High-Speed Modem Setup Setup for Mariner High-Speed Telemetry Simulator (PDP-4) Simulation of High-Speed Telemetry Data Setup of the Tracking Data Simulator (PDP-4) Setup for the TTY Encoder Direction and Coordination During a Mission- Data Control (DACON) AFETR Data: Computer Processing Support Group Duties for Mariner Cruise Phase Bulk Data Handling and Distribution b. Engineering Project Documents No. Title EPD-122, Addendum I EPD-122, Revision 2 EPD-122, Revision 1 EPD-122 EPD-173, Revision 1 EPD-173 EPD-203 Space Flight Operations Plan, Mariner Mars 1964 Space Flight Operations Plan, Mariner Mars 1964 Space Flight Operations Plan, Mariner Mars 1964 Space Flight Operations Plan, Mariner C: Missions P-70 and P-71 Space Flight Operations Test Plan, Mariner C Space Flight Operations Test Plan, Mariner C Space Flight Operations Facility Capabilities Document for Mariner C 6/25/65 10/12/64 10/12/64 10/12/64 10/12/64 10/12/64 10/12/64 10/12/64 10/28/64 7/12/65 10/12/64 Date 7/5/65 10/28/64 8/17/64 7/15/63 5/11/64 8/1/63 3/2/64 3O JPL TECHNICAL REPORT NO. 32-881 Table6. (Cont'd) c. Manuals Title Date Mariner Data Processing Manual, Revision 1 Mariner Data Processing Manual Mariner IV Progress Report No. 4 (Launch -t- 155 Days to Lcunch + !_n5 D__y._) Mariner IV' Progress Report No. 3 (Launch + 109 Days to Launch + 154 Days) Mariner IV Progress Report No. 2 (Launch + 79 Days to Launch + 108 Days) Mariner IV Progress Report No. 1 (Launch to Launch + 80 Days) Mariner Mars 1964 Quarterly Report for the DSN/SFOF Mariner Mars 1964 Quarterly Report for the DSN (SFOF) 4/15/65 10/13/64 7/5/65 6/7/65 4/15/65 3/29/65 5/17/65 2/15/65 Space Flight Operations System Design Specifications, Mariner C Mariner (111)Mars 1964 Technical Bulletin No. 1 Mariner (IV) Mars 1964 Technical Bulletin Nos. 1 through 307 Original 3/12/64 11/9/64 11/30/64 through 10/4/65 _To consolidate the planning and preparation for the mission following maneuver, Encounter Planning Groups were organized and defined. Regular sessions were held and priorities of effort were established in these meetings. From the encounter planning activities, test require- ments were established, using the SAF/PTM/SFO facili- ties. During the operational readiness tests there was a SAF/PTM/SFO/DSIF participation. The appointment of a specific Communications Project Engineer for Mariner greatly improved the interface between SFO and Communications. This improved inter- face accomplished the encounter ground communication requirement of Fig. 35 with very few problems. An SFOF Project Engineer supported the Mariner 1V mission from prelaunch through mission termination. Provisions supported by this function were: (1) continu- ous facility hardware status, and (2) implementation of hardware for both flight operations and test activities. C. Encounter Readiness Tests In preparation for encounter, a number of tests were scheduled involving nominal encounter operations and possible encounter operations in a failure mode condition (Table 7). These tests exercised the analysis and opera- tions personnel in the performance of their duties and met the objective of developing and maintaining the pro- ficiency required to participate properly during the planetary encounter portion of the mission. Table 7. Encounter readiness tests J Date 4/13/65 4/15/65 4/20/65 5/4/65 5/4/65 5/27/65 6/3--4/65 6/23-24/65 6/24.-25/65 6/24/65 6/25/65 6/28-30/65 7/6-7/65 7/8-9/65 7/12/65 Description DC-18 Test DC-18 Test Nominal SPAC encounter Roll position slmulation Data Mode 4/1 transmission from DSIF 51 Encounter backup mode Preliminary operational readiness Command and occultation Command and occultation Encounter alternate plan Encounter alternate plan Operational readiness Operational readiness Operational readiness Operational readiness 1. Roll Increment Tests The first test in preparation for encounter involved the use of De-18 (gyros on inertial control: positive roll increment). The test was performed on 13 April 1965 and was conducted from the Mission Support Area (MSA) using the PTM as the data source. The objectives of the test were: 1. To familiarize personnel with an encounter proce- dure that uses De-18 as a backup to observe degra- dation in spacecraft parameters 2. To familiarize the Spacecraft Performance Analysis and Command (SPAC) and Space Sciences Analy- sis and Command (SSAC) personnel with the space- craft while operating in a simulated radiation field of Mars 3. To familiarize personnel with the communications time at encounter, and to perform operations with this constraint 31 JPL TECHNICAL REPORT NO. 32-881 . I °o0 I o I °00 | ! 0 w o_ 32 JPL TECHNICAL REPORT NO. 32-881 4. General DC-18 familiarization with roll position determination from the down link signal 5. To familiarize personnel with the procedure to be followed in going from a DC-18 mode to the DC-15 mode of operation A second DC-18 test was conducted from the MSA on 15 April 1965 using the PTM as the data source. This test differed from the first DC-18 test in that the DC-18 was used as the primary mode of operation. A roll position simulation test was performed using DSIF 11 and the SFOF on 4 May 1965 between 2210 GMT and 0830 GMT. The test was conducted to deter- mine the effect of spacecraft roll on the ground received signal. 2. Scan Platform Positioning In addition to positioning the scan platform during all the tests involving an encounter sequence, several indi- vidual scan positioning tests were conducted. All plat- form positioning tests involved the PTM with the exception of one which used simulated data transmitted from DSIF 51. These tests were proficiency tests for the science operations personnel involved in positioning the scan platform via ground command. 3. Procedure Verification Tests A preliminary operational readiness test was conducted on 3 and 4 June 1965. Simulation and Mariner PTM data were used for this test. The objective was to verify the adequacy of encounter procedures and to verify the smooth operation of all of the interfaces between the var- ious components of the DSN. A nominal SPAC encounter test was performed on 9_ April 1965 using the Mariner PTM as the data source for telemetry and simulating DSIF data and functions. The test objectives were (1) to determine the feasibility of the nominal encounter sequence in terms of command response, timing, and opportunity for thorough evalua- tion of data; and (2) to offer Mariner operations person- nel, SPAC, and SSAC an opportunity to familiarize themselves with the proposed encounter sequence. A tape containing Mode 4/1 data was sent to DSIF Station 51, and on 4 May 1965 it was transmitted to the SFOF. The data transmission occurred during those hours that the actual encounter data would be transmit- ted from Johannesburg. The purpose of this test was to simulate transmission conditions that were closely related to those occunJng during the actual picture playback from Station 51. 4. Backup Mode Test An encounter backup mode test was conducted, using the Mariner PTM as data source on 27 May 1965 from 1500 to 0000 GMT. The backup mode test measured the capability of the SPAC and SSAC operations personnel to respond adequately to failures and/or anomalies oc- curring during the encounter phase of the Mariner n_- siom The SAF/ffFM Test Team, at their discretion, induced failures and anomalies for which it was SPAC's responsibility to determine the nature of the problem and make recommendations for corrective action. 5. Tests from Flight Path Analysis Area Two command and occultation tests were run from the Flight Path Analysis Area (FPAA) at times shown in Table 8. Table 8. Command and occultation tests Date Start End Time (GMT) X23 June 1965 24 June 1965 24 June 1965 25 June 1965 X 2000 0445 2000 0445 There were four principal objectives of the occultation test: . Checkout the ability of the orbit determination pro- gram (ODP) to provide predictions with the re- quired accuracy in Mode 2 and to test the plotting capabilities of the SFOF Data System . To verify the ability of the data processing system of Goldstone to provide computations of occultation data and open-loop receiver synthesizer settings . To train the station personnel in the operation of the open-loop receiver and tape recorder system, as well as in the special handling of occultation dopp- ler data . To establish procedures for rapid reacquisition of two-way communications in emergency conditions, using the 100kw Station 13 transmitter 33 JPL TECHNICAL REPORT NO. 32-881 Two encounter alternate plan tests were conducted on the following dates: Date (GMT) Time 24 June 1965 1500-2300 25 June 1965 1500-2300 The encounter alternate plan tests were basically logic- diagram checkout tests used to establish and verify a set of telemetry criteria that detcrmined the logical state of the spacecraft; they also served as a basis for decision concerning nonstandard ground command action and deviations from the nominal encounter sequence plan. 6. Operational Readiness Tests The first operational readiness test was conducted from 0700, 28 June 1965 GMT to 0400, 30 June 1965 GMT. The objective of the operational readiness test was to provide verification of the operational readiness of the combined DSN prior to the Mariner IV planetary en- counter. The test also verified the adequacy of encounter procedures and the smooth operation of all the interfaces between the various components of the DSN. The final operational readiness test was conducted from 0700, 6 July 1965 GMT to 2230, 7 July 1965 GMT. The objective of this test was to verify the operational status of all equipments, procedures, techniques, and pro- grams required during the Mariner IV encounter. An additional operational readiness test was conducted from 0100, 8 July 1965 GMT to 1300, 9 July 1965 GMT. The objective of this test was to verify that the Analysis Areas and the DSN were fully prepared to commit to support the Mariner IV encounter. On 12 July 1965 a final operational readiness was con- ducted. Previous tests had simulated many anomalous conditions, and the purpose of this test was to perform an encounter sequence with no anomalous events. 7. Command-Loop Lockup A number of command-loop lockup exercises were conducted, starting in April, 1965 with Station 13. These exercises were conducted to maintain a high degree of proficiency at all stations involved in the planetary en- counter of Mariner IV. Table 9 provides tabular infor- mation of the command-loop lockup exercises. Data 8 April 28 April 28 April 30 April 12 May 1 June 3 June 16 June 28 June 28 June 30 June 1 July 6 July 6 July Table 9. Command-loop Iockup exercises Number of Station times 11/13 1 51 1 11 I 42 2 11/13 1 11/13 13 11 51 11 51 11 51 11 1 7 July 13 1 Remarks Station 13 (lOOkw) transmitter Station 13 (lOOkw) transmitter, Station 11 receiver Station 13 (lOOkw) transmitter Station 13 (100kw) transmitter 1 1 Station 13 (lOOkw) transmitter 2 Once with each RWV Using 10kw transmitter 34 JPL TECHNICAL REPORT NO. 32-881 IV. POST-LAUNCH TESTING A. Life Test Program A life test on the subsystem level was a Mariner C Project requirement. In carrying out this directive, lists of equipment that should be subjected to life testing were compiled, and members of the Environmental Require- ments group worked with the vaiSous _.._2.nt pe._on- nel to delineate test philosophy and establish general rules for testing the equipment. Cognizant personnel outlined minimum test plans for each unit-for the operating modes, measurements to be made, instructions for test performance, acceptable performance, and a format for a monthly status report. Individual test plans were completed prior to 30 November 1964. Monthly status and progress reports were submitted to the Environmental Requirements group-where they were integrated and, between February and July 1965, the status of life testing was published each month. Ab- normal results were reported on detection. On 3 June, the Spacecraft System Manager's Office announced the decision to terminate all life testing by 30 July 1965, and final reports on the individual life tests were called for by 16 August. Subsequently, a decision was made, jointly with the Voyager Project Office, to continue testing on the radio and on the Canopus tracker. A summary of life tests is given in Table 10. Test item Ref. Cosmic Dust Detector Ionization Chamber (Cosmic Ray Telescope, not tested) Trapped Radiation Detector Magnetometer, operated at flight temperatures ISolar Plasma Probe Data automation system Planetary scan Television Data encoder Command 24A1 26A1 21A1 25A1 33A1 33A2 33A3 32A1 32A2 32A3 32A4 20A1 20A2 20A3 20A4 20A5 31A1 31A2 31A3 C36 6A1-13 3A1_ Table 10. Summary of life test status Serial No. TA 111 TA TA 2 O MC-0 } MC-0 MC-0 MC-0 2 MC-7 TAW MC-7 3 TA TA MC-TAM 3 2 Life testspan Start End 2/25/65 9/15/65 10/16/64 7/30/65 6/18/65 7/30/65 10/12/64 7/30/65 (Final report 12/22/64 not received) 12/23/64 1 2/12/65 1/23/64 7/30/65 /12/ 5 12/23/64 8/24/64 7/4/65 8/4/64 7/30/65 7/10/_ 4/26/65 7/9/64 7/31/65 Regular reporting, date of start 2/28/65 2125/65 2/25/65 2126/65 2/23/65 2/24/65 2125165 2/22/65 3/29/65 i V_¢. Test environment Temperature Low High Cycling Observed effects Failures Degmd_ tlon x No data x X 35 JPL TECHNICAL REPORTNO. 32-881 Test item Ref. Serial No. Radio 2 Tape machine 16A1-5 Power 4A Baffery" 4A14 Attitude control electronics 7A1 Sun sensors 7PS6 Earth detector 7ED6 Canopus sensor 7CS8 Narrow-Angle Mars Gate 7MG1 Solar pressure control assy. 7PC Jet vane actuators 7JV Attitude control gyros 7A2 Central computer and 5A sequencer Scan actuator 31A4 Dampers, retarder and 8M1 timer Science cover latch actuator 31M2 Pyro control 8A1 8A2 Post-injection propulsion system • Degradation was normal aging process. Table 10. (Cont'd) MC-0 2 02 6 7 8 9 18 002 2 2 004 011 8 Life test span Start End 1/27/65 7/30/65 8/4/64 2/23/65 8/17/64 7/16/65 2/11/64 11/2/64 2/11/64 4/27/64 8/19/64 5/3/65 5/1/64 12/18/64 11/28/64 7/28/65 6/15/64 8/30/65 6/15/64 5/3/65 5/12/64 6/18/65 12/18/64 7/31/65 5/12/64 4/19/65 10/5/64 8/30/65 2,4,7,32 012 MC-TA Clll 1 C126 C104 C132 C4 C-Proto 1003 t4 OO3 8/24/64 4/1/64 411164 1119164 8/31/64 2/10/64 8/30/65 8/6/65 8/15/65 5/5/65 7/30/65 12/30/65 Test environment Regular reporting, date of Temperature start Vac. Low High Cycling 2/28/65 x -- x -- X -- X -- 2/16/65 x -- -- x I X X -- -- 2/16/65 x -- x -- X _ X -- X -- X -- 3/3/65 -- -- x -- 1/4/65 .... 2/2/65 .... x m -- -- 2/I/65 .... 3/1/65 -- -- x -- 3/1/65 -- -- x -- 3/1/65 x -- x -- 2/17/65 -- -- -- x 2/11/65 x1/29/65 Observed effects Failures Degrada- tion x x (some set observed from long storage) X -- -- X X X -- X X X -- X -- X X -- B X m m 36 1. SignificantFindings Someunits survived during the test period with no apparent degradation; others showed degradation and/or failure. The varying degree to which units were affected is summarized here. a. Units showing neither degradation nor failure. Umtte_ed • • t-J z°xCCU_ operation time, hr Cosmic Dust Detector 5,514 Trapped Radiation Detector 7,033 Data automation system (DAS) 4,800 Planetary scan 3,194 Data encoder 9,000 Attitude control (A/C) electronics 11,000 Earth detector 9,268 Solar vane actuator 5,846 Thrust vector control 7,828 Pyro control 8,256 b. Units showing degradat/on. There were six units that exhibited some form of degradation during life test. The reduced performance and probable causes are described. Magnetometer. Degradation in the sensor was detected early in the life test; it was probably due to a defective lamp and cell that were installed prior to life test and were not flight acceptance (FA) or type approval (TA) tested. Implications of this degradation are not dear. Vacuum operation of the unit disclosed a leak in the hermetic seal of the infrared detector, which is part of the sensor; this fact supports the argument for perform- ing the life test in vacuum. Television. Slight degradation in picture quality was detected; this could be the result of overexposure caused by shutter failure on several occasions. The shutters are known to be limited life items and were not part of this life test. (A longer life test might have been helpful in evaluating the causes of this degradation-) JPL TECHNICAL REPORT NO. 32-881 Command subsystem. No failures were detected in 8000 hr of life test; the total accumulated time on the life test unit, including other tests, was 10,800 hr. Only the voltage controlled oscillator (VCO) frequency exhibited some drift. The drift was consistent and would indicate that eventually the parameter would drift out of spec. It is difficult to estimate precisely the time when this amount of drift would occur--the cause of which has not been investigated; a longer life test might have been help- ful in evaluathtg _1_.......UItlu U_. U_lt..,. Video storage. On 3 February 1965, after 6 mo of non- operation, power was applied to the video unit. A record sequence was initiated, and the run-up and run-down times of the motor were measured. The run-up times decreased from 2.15 see on the first picture to 1.95 see on the seventh and succeeding pictures. The run-down time increased from 4.65 see to 5 sec after seven pictures. The conclusion is that the unit did take a set during the period of vacuum storage; since it was hermetically sealed and did not leak during test, the degradation must have been caused by age, not vacuum. How serious a set would exist after longer periods of storage could not be evaluated on the basis of this life test because of the test interruption to facilitate a decision on the Mariner 1V cover drop exercise. Narrow-angle Mars gate. Some degradation in sensi- tivity was observed during the first 20 hr of life test; however, the unit stabilized and operated without failure or further degradation for the remainder of the life test (total operating time, 8210 hr). Attitude control gyros. Gyro motor bearing wearout was observed after about 5000 to 5500 hr. As expected, the high operating temperatures encountered during the 1000 hr vacuum run seem to have accelerated this wear- out. Lubrication breakdown and subsequent bearing wear is always accelerated by high operating tempera- tures. It is suspected that the bearing preload was too high on these Mariner C gyros; meeting the requirement on the anisoelastic coefficient may have been the cause of the sacrifice of some bearing life. The lifetime of the gyros was demonstrated by this test to more than exceed the mission requirements, although in general, the life test gyros' performance was somewhat poorer than was expected. The additional handling involved with this unit may account for some of this degraded performance dur- ing the life test. c. Units showing degradation and failure. The follow- ing units exhibited either sudden failure or degradation leading to failure. 37 JPL TECHNICAL REPORT NO. 32-881 Radio. The life test, begun on 30 January 1965, is con- tinning on this unit. Some degradation is being observed in the oscillator circuit but, since the test is continuing, further analysis is being postponed. The redundant ex- citer failed after 262 hr on life test; however, this is not considered life-test failure, since this unit's performance had degraded considerably prior to life test. Battery. One failure, resulting from a manufacturing defect, was detected early enough in the project to result in a modified FA test procedure. Some qualitative infor- mation was obtained on the battery's degradation rate as a function of temperature. Cell voltage was found to be an adequate parameter for the detection of battery failures. Degradation was detected on all five batteries life tested. One other failure was observed during the life test, but this was considered a normal end-of-life for a battery operating at an elevated temperature. The battery is the only case observed during the life test when degradation in unit's operation parameters led to a complete failure. The degradation observed on the gyros may have been another example, but this test was interrupted before the failure became catastrophic. d. Units experiencing _ailure. Four units experienced sudden failures during the life test without earlier detec- tion of degradation. Ionization Chamber. The same failure mode experi- enced by the flight unit was duplicated in the life-test unit by stimulating it to the high counting rates experi- enced in flight. Although this was actually beyond the scope of the test, it illustrates the use of life-test hardware to support the mission. No failure and/or degradation was detected that was not the result of the induced failure. Plasma Probe. Ort 24 December 1964, after a total of about 1163 hr, the R-342 resistor failed. This type of failure is identical to the one that occurred in the Mariner IV instrument during flight. A life test per- formed early enough, under operating conditions similar to those encountered in flight, would have revealed the design deficiency that caused this failure. Power Unit. The life test was terminated after 8300 hr because the 3-phase inverter drew excessive power from the maneuver booster regulator. The input filter capaci- tor of the 3-phase unit was short-circuited as a result of electrolyte leakage through the terminal seals of the ca- pacitor. This failure mode might have been obscured if the life test were conducted at ambient laboratory condi- tions (the unit was tested in vacuum). When the unit was removed from life test, another failure was detected. This was a non-catastrophic failure and was not detected dur- ing the life test. The cause of failure could not be dis- covered but some mishandling (during manufacture) of the failed part (2N1016D transistor) is suspect. Central computer and sequencer (CC&S). All life test failures observed appear to be mechanical in nature and may be the result of TA testing. Briefly, these failures consisted of two corroded jumper wires, two corroded component leads, one cracked diode, one broken unit of prepackaged magnetic wires. Outside of these, the unit performed without any failure or degradation for 10,000 hr. It should be pointed out that the temperature cycling during the life test may have been responsible for such failures as those detected in the prepackaged magnetic wires, as well as the cracked diode. However, since the unit was subjected to a severe TA test which may have contributed to these failures also, the actual cause cannot be determined. e. Ground Support Equipment Problems. Canopus sensor. Sole failure reported against the life test sensor resulted from damage caused as a result of ground support equipment (GSE) failure. Unit accumu- lated 5831 hr of operation. Sun sensors. No failures and/or degradation was de- tected after 7765 hr of operation. The test equipment was less accurate and stable than the sensors, causing faulty parameter drift information. Mechanical devices. No problems with the actual hard- ware were encountered during the test. Some problems with the GSE were experienced. The science cover sole- noid and the scan actuator operated without any detect- able failure or degradation. The long term oil leakage test performed on the dampers was not very successful because of the problems with the GSE. However, the measured weight losses indicated that the outgassing was in such small amounts that any effect on spacecraft per- formance would be negligible. 38 92. Summary of Significant Interruptions a. Building power failures. The following subsystems reported several interruptions to the life tests resulting from building power failures: Data encoder Command subsystem Radio Power subsystem Attitude control electronics Sun sensors Earth detector Canopus sensor Narrow-angle Mars gate Attitude control gyros and gyro electronics b. Thermalvacuum system problems. The life tests of the following units were interrupted as a result of a thermal/vacuum system problem: Magnetometer. On 27 and 28 January 1965, repairs were performed on the vacuum chamber, resulting in a 36-hr interruption to the life test. (This was not a very severe interruption since, at the beginning of each week, the unit was removed from the vacuum chamber to check the lamp and cell ignition voltage of the magne- tometer sensor.) Plasma Probe. Test was interrupted on 1 January 1965 because of vacuum problems. Life test was not resumed until 20 February 1965. Radio. The pump system in the vacuum chamber failed on 25 February 1965. Test was interrupted until 1 March 1965. Power subsystem. On 5 June 1965, the temperature control system of the vacuum chamber failed, causing the temperature of the unit to reach 250"F at one tempera- ture monitoring point (see P/FR 7825 for details). Scan actuator and support; dampers, retarder and timer; science cover latch actuator. One of the major objectives of this particular life test was the detection of slow oil leaks from the hydraulic devices. To accomplish this, small condensing mirrors were placed in strategic JPL TECHNICAL REPORT NO. 32-881 points throughout the chamber. These condensing mir- rors could be used to detect traces of oil by measuring the changes in their optical properties. However, the master valve controlling the flow of liquid nitrogen to the entire Environmental Test Laboratory, where the life test on this unit was performed, was accidentally closed, resulting in backstreaming of the condensible material from the cold trap into the chamber and contaminating the mirrors. c. Life test operations support equipment (OSE) problems. Sun sensors. Several days of life tests were lost due to abnormalities in solar simulator operation and building power shutoffs. Canopus sensor. On 12 April 1965, problems with OSE power supply caused a transformer in the tracker to burn out. Narrow-Angle Mars Gate. Planet simulator power sup ply failed and had to be replaced. d. Test interruptions for mission support activities. Ionization Chamber. Special tests were performed to investigate the flight failure. A/C electronics. At one point in the life test, the unit was withdrawn for use in the investigation of a flight spacecraft problem. Tape machine. The life test of the tape machine, which consisted of storage in a controlled thermal/vacuum envi- ronment, was terminated prematurely to investigate pos- sible problem areas which may have been encountered by the Mariner IV unit. e. Interruption for radiation tests. Life tests on the planetary scan, the television, and the Narrow-Angle Mars Gate units were interrupted to allow special radia- tion tests to be performed. f. Interruptions from various causes. DAS. The DAS was initially placed on life test 23 December 1964 using two MC-1 real-time (RT) trays (20A2 and 20A4). After 990 hr of life test, the MC-1 RT failed and was removed for evaluation. The MC-7 (TA) RT was installed in place of the failed unit and the test was resumed on 14 February 1965. 39 JPL TECHNICAL REPORT NO. 32-881 Television. From 4 September 1964 to 8 September 1964, there was a shutter failure. From 12 September 1964 to 21 September 1964, a shortage of personnel to conduct life test caused interruption of testing. Battery. All life tests prior to August 1964 were con- ducted at atmospheric pressures because of the unavail- ability of vacuum facilities. From 16 August 1964 through 19 August 1964, life tests on batteries 6 and 9 were inter- rupted to place units in vacuum chambers. A/C gyros and gyro electronics. The A/C gyros and associated electronics underwent a series of tests which were considered as phases of the life test. Going from one phase to the other resulted in some short interruptions. g. Life tests reporting no interruptions. Cosmic Dust Detector Cosmic Ray Telescope Trapped Radiation Detector Solar pressure control assembly Jet vane actuator CC&S Pyrotechnic control subsystem Post-injection propulsion system (PIPS) 3. Recommendations for Future Programs a. Central test facility. An examination of the sum- mary of significant interruptions to the various life tests reveals that most of these were the result of either build- ing power shutdown or test facility failure. Indeed, the objectives of one life test (mechanical devices) were not fully realized because of a facility failure. These facts would suggest that future life tests be performed under more controlled test conditions with stringent safety con- trois and constant, transient-free, power sources. An efficient solution to this problem would be to pro- vide a central test facility for all life tests, which would give the following advantages: 1. Only one constant-power, transient-free source would have to be provided. 2. Economy, with respect to the number of trained test personnel necessary to operate the facility, would result. 3. Supervision and rigid control of the necessary safety provisions would be provided. 4. A central location for test equipment having long mean time between failures, as well as the special- ized facilities necessary for their maintenance, would be provided. 5. Interference of life tests with other test activities that might be going on at the same time would be prevented. Almost mandatory to an efficient operation of such a facility would be the provision of automatic test check- out equipment for the hardware on test. This not only would minimize the test personnel necessary to staff the central test facility but, also, would minimize the danger from mishandling and operator error during the equip- ment checkout procedures. Continuous monitoring of the hardware on test, either by technician or automatic equipment, is always desir- able in detecting and analyzing intermittent and/or periodic failures. Automated checkout equipment is the more efficient and, in most cases, the more economical. b. Central data storage and analysis. A central data storage and analysis area would contain all test histories for an entire test program and would provide easy access to all needing such data. Such a library would facilitate the comparison of related effects and the analysis of these effects. c. Selection of life test units. To provide data on the flight units, the life test units should resemble the flight units as closely as possible. Any incidents in the history of life test units that would not normally be experienced by the flight units, and the effects of which on the life test units cannot be fully evaluated, should be avoided. The use of TA units, which have undergone a wide range of tests more severe than the FA tests, is not ideal be- cause the effeet of the TA environment on the life ex- pectancy of the test items cannot be fully evaluated. For example, the observed cracked diode in the CC&S may have been the result of the temperature cycling during life test or may have been the result of TA testing. The use of more than one unit for the life tests would be a great aid in the analysis of the effects observed during life tests and would, also, provide a better sample from a statistical viewpoint. Effects peculiar to a spe- cific unit, but not typical of the hardware design, can 4O JPL TECHNICAL REPORT NO. 32-881 "be eliminated more easily. On the other hand, time- dependent effects that are typical of the flight equipment will become easily recognizable when exhibited by more than one test unit. d. Use of thermal/vacuum facilities. The desirability of using thermal/vacuum facilities has been demon- strated during the Mariner C life tests by the significant number of failures and degradation which would not _fft--ve been .1.... _Aa _ _1_,_ lff_ t_t_ were verformed at ambient laboratory conditions. Good examples of these failures and/or degradation are the ones occurring on the Magnetometer, the power subsystem, battery, and the A/C gyros. e. Desirability of performing life tests in space simu- lated environments. In general, it is desirable to minimize the number of unknown effects contributing to the unit's degradation and/or failure. Since it is important to evalu- ate the hardware's ability to perform its mission, and since insufficient data were obtained to validate accel- eration techniques, it is most desirable to duplicate, as much as possible, the mission environment. Accelerated testing may be a valuable tool ff enough is known about the acceleration process. If the acceleration process is not properly understood, more may be ob- scured by an accelerated life test than revealed by it. However, when only short periods of time, relative to mission duration, are available for life testing, the use of accelerated techniques may be unavoidable; but the life test in such a circumstance becomes simply a tool for the generation of hardware failures in the hope that signifi- cant design deficiencies would thus be revealed. f. Radiation environment. The experience with the ion chamber suggests that it may be possible and desirable to simulate a radiation environment which will be expe- rienced during the mission and to subject test units to it. Life test may be the proper place in which to evaluate the effects on hardware of prolonged exposure to radiation. g. Digital electronics. To aid in the detection of degra- dation and the analysis of failures, some provisions should be made on digital electronics to perform 'param- eter" measurements on vital test points throughout the unit. h. Life test scheduling. Life tests should be performed early enough so that results may be implemented in the final hardware design. 4. Conclusions Reporting on failures of life test units has been con- fined here to those failures resulting from inherent equip- ment weaknesses that, in ham, were caused either by improper manufacturing procedures or by the use of defective components. Overall test objectives of the life test program were achieved: that is, (1) the analysis of such performance effects as degradn_Oon a_ a function of time, (2) the iden- tification of critical failure modes and influences of space environments, (3) the support of inflight problem diag- nosis, and (4) the provision of test information indicating the nature and extent of flight hardware deterioration. In addition, there has been an accumulation of experi- ence and useful data that will aid in conducting future tests. B. Electron Radiation Tests The effects of possible radiation hazards to be encoun- tered by Mariner IV at the time of encounter with the planet were investigated. Tests _ were conducted on several Mariner C electronic subsystems with a Dynamitron accelerator at fluxes in the range 1 x 10 r to 2 × 1011 eleetrons/em2-sec, and energy levels of 0.5 to 2.7 Mev electron energy. Aluminum or copper foils were placed in the beam path upstream from the test item to scatter the beam uni- formly (±20%) over the bombarded subsystem surface. Each subsystem exhibited a repeatable circuitry failure threshold. The failure thresholds were rather insensitive to either the electron energy or integrated flux, but they were very dependent on the flux rate; Failure thresholds for the subsystems tested ranged from 8 × 10 s to 6 × 109 electrons/cm2-sec. One subsystem was found to operate satisfactorily in a flux as high as 5 × 101° electrous/cm2-sec. JPL became concerned with the effects of electron radiation on Mariner IV subsystems when R. Davies, a radio astronomer from Jodrell Bank, England, reported that his observations of radio noise at 21 cm showed that Mars had a brightness temperature of l140*K. Since the infrared (IR) temperature of Mars is 250"K, this sug- gested that Mars had radiation belts several times as strong as those of the Earth. aTests were conducted at C,eneral Dynamics Convair, San Diego. 41 JPL TECHNICAL REPORT NO. 32-881 Consequently, an electron testing program was begun on those Mariner IV subsystems that were judged to be most susceptible to radiation, that were critical to the mission, and that might be ground controlled in such a manner as to avert disaster ff their radiation sensitivities were known. The systems chosen for testing were: (1) the Canopus star sensor, which is a sensitive electro-optical instrument containing a special photomultiplier tube; (2) the data encoder, a complex electronic switchyard that prepares the scientific and engineering data in proper format and sequence for telemetry; (3) the plane- tary scan platform, a wide-angle optical sensor whose purpose is to point the narrow-angle television camera lens at Mars; (4) the television system; (5) the narrow- angle Mars gate (NAMG), a narrow-angle Mars detector used as a back-up start for the television camera; and (6) the gyro and attitude control system. 1. Environmental Estimates a. Earth's Van Allen belts. The electron flux in the outer Van Allen belt at approximately four Earth radii was measured by Explorer XIV during the 5 mo from October 1962 through February 1963 (Ref. 3). Typical omnidirectional fluxes in the heart of this zone were: Jo (E > 40 key) = 3 X 10' electrons/cm2-see Jo (E > 230 key) = 3 x 108 electrons/cm2-sec Jo (E > 1.6 Mev) = 3 X 105 electrons/cm2-sec Temporal variations were observed to raise and lower the above values by as much as a factor of 10. b. Estimated radiation belts at Mars. With the as- sumptions that Mars had a magnetic field and that the electron flux on Earth and Mars was equal at points of the same field strength, rough estimates of the radiation belts of Mars were made based on assumptions of strength of the Mars magnetic dipole momer_t. Since at the begin- ning of the investigation the moment of Mars was not known, early estimates of the Mars radiation belts were highly conjectural. If the Martian magnetic moment were less than that of the Earth, then it was expected that the spacecraft would encounter fluxes less than 8 X 10 r electron/cm2-sec at Mars. H. R. Anderson made an esti- mate based on various assumed MartiarL magnetic mo- ments; the containment power of the magnetic field gives extreme upper limits of electron fluxes through which Mariner IV would fly, as follows: Condition Max. omnidirectional flux (E, > 2 Mev) 1.2 X 109 electrons/cm2-sec 1.2 X 1011 electrons/cm2-sec Since Mars was expected to have a magnetic moment < 10% of Earth's (M_ < 0.1 M_), it was considered highly unlikely that the spacecraft would encounter TEST ITEM --_ \ /- STAINLESS STEEL / VACUUM CHAMBER MOVABLE CARBON _ _ J BEAM DEFINING / \ _¥ S.OTTERS / \ \ ELECTRON BEAM \ / \ \ FROM _ / _ F'---I \ ACCELERATOR| _ _J _I I I _L I_ ...._I] ....._...--_" .L I TO LIQUID N z TRAP t -'r-----P'l _ _ .I _ i I I _ AND VACUUM PUMPS BEAM STOP AND ......... _ / X_FARADAY CUP FARADAY CUP _'_SCATTERING _ulLb _ _ ...... VIEWING PORT Fig. 36. Electron radiation test setup 42 electron fluxes as large as 1.2 × 109 electrous/cm_-sec; rather, they were expected to be < 3 × 107 electrons/ cm2-sec. 2. Test Activity ao Apparatus. Figure 86 shows the geometry of the test setup. The electron beam from the accelerator may be stopped by the pneumatically operated Faraday cup, allowing the beam current to be preset to the requiaed test value. The scattering foils are 24 in. apart; their function is to broaden the beam spatially to cover the test item with a uniform electron flux. The four carbon shutters could be moved into, or out of, the beam to achieve a rough control over the beam intensity at the test surfaces. During the test, vacuum in the chamber was maintained at 10 -5 torr or below. The shape of the beam after scattering was determined by driving a Faraday cup across the vacuum chamber with a screw mechanism and monitoring the collected current with an electrometer driving a strip chart re- corder. Several flux maps obtained in this way are shown in Fig. 37, together with roughly fit Gaussian curves. The flux distribution was shown to be symmetrical about the beam centerline by traverses below center, hori- zontally, and at 45 deg from vertical. The curves were taken with one foil only in a traverse across the center of the vacuum chamber 26 in. from the foil. Similar curves were obtained by using two foils as shown in 1.0 0.13 % % 0.6 cr" 0.41 bd 0.2 • o OO e: C) 06-Mev ALUMINUM -- [] 23-Mev COPPER A I.O-Mev ALUMINUM 0 12 I0 8 6 4 ABOVE BEAM CENTERLINE,in. Hg. 37. Electron flux distribution in the area of the test flux 2 0 JPL TECHNICAL REPORT NO. 32-881 Fig. 86. It was found that changing the position of the carbon shutters did not change the curve shape, within experimental accuracy. It is apparent that the single aluminum (Z = 13) foils are useful electron scatterers for this test only below 1 Mev. Accordingly, two copper foils (Z = 29) were used for the higher energies. The lower Z foils were used when- ever possible because the bremsstrahlung or X-ray pro- duction, as well as the scattering ability, ri_es as Z" (Ref. 4). With the test geometry described above, directly measurable beam fluxes at the subsystem surfaces varied from 107 electrous/cm2-see to a maximum of 2 × 1011 electrons/cm2-sec. The low fluxes were attain- able only at lower electron energies where corona dis- charge inside the accelerator beam tube was not excessive. The upper flux limit is determined by accelera- tor stability at low energies and by the beam current that will bum a hole through the foils at higher energies. Electron energies from 0.5 to 2.6 Mev were used in the test. Flux monitoring was achieved by using a Faraday cup mounted adjacent to or, ff possible, on the front surface of the test item. The flux measurements are estimated to be accurate to within -+-20%. Electron energies quoted in the test results are corrected for en- ergy loss due to foil traversal. Taking into account the energy spread due to foil-induced straggling and the ac- curacy of the measured accelerator voltage, the electron energies are accurate to roughly 10%. The subsystems were mounted in their magnesium flight cases with the beam incident on the flight case. At the thinnest portions of the case, the electrons had to penetrate about 0.6 g/ern 2 of magnesium, with loss of about 0.8 Mev. Note that the Canopus sensor, planetary scan preamplifier and motor, television preamplifier and optics, and the NAMG are not protected by a flight ease. A nominal 0.2-Mev electron energy loss occurs in pene- IxalJng the respective aluminum shields. The test items were mounted inside the vacuum chamber on an adjustable aluminum pedestal and plat- form. With the exception of the planetary scan, television, and NAMG subsystems, the test items were located 39 in. from the last scattering foil. The above noted exceptions were mounted in the chamber center 26 in. from the foil, so as to utilize external optical stimuli. The subsystems were connected by 35-ft cables to the OSE located outside the accelerator target area. JPL TECHNICAL REPORT NO. 32-881 b. Test procedure. Following exploratory test runs on the Canopus sensor to determine general areas of mal- function, the procedure was to initially irradiate the sub- system at 0.56 Mev for 20 min at a flux of 108 electrons/ cm2-sec. This was chosen as a flux/energy combination yielding a conservative simulation of radiation levels that the spacecraft might experience during the encounter sequence. The program was then to increase fluxes by approximately half-order-of-magnitude steps (pausing at each level for subsystem checkout by the cognizant en- gineer), until a malfunction ensued or maximum accelera- tor flux was reached. The electron energy was then increased and the flux values again stepped through. Whenever a malfunction occurred, the beam flux was adjusted downward in small steps in order to ascertain a threshold flux level for malfunction. The subsystem was also monitored after each malfunction to determine if permanent damage had occurred. Monitoring was done by moving the retractable beam stop in to intercept the beam. Subsystem recovery was then evidence that RF fields, power-line loading or other phenomena associated with operating the accelerator were not responsible for the observed behavior. 3. Results Figure 38 is a plot of the malfunction thresholds ob- served for the various subsystems. The various modes of malfunction are presented in Table 11. _n % z o fY w w x" I0111 i010 - 109 iO 8 IO ? 0 (_) PLANETARY SCAN PLATFORM (_ TELEVISION ® NARROW-ANGLE MARS GATE I (_) TESTED LIMIT FOR GYRO & ATTITUDE I' CONTRiL 0,5 t,O 15 2lO 215 3 lO ELECTRON ENERGY, Mev Fig. 38. Radiation thresholds of malfunction for various subsystems in test Subsystem Canopus sensor Data encoder Planetary scan Television Narrow-Angle Mars Gate Gyros and attitude control Table 11. Subsystem failure modes Radiation-induced failure mode Decrease in roll error saturation voltage below ___2.0 v Null voltage offset Apparent star intensity increase with no star in view Commutator sync loss predominant at energies below 1 Mev Demodulation lock loss above ! Mev Event counter change of state Switch from Mode 1 to Mode 4 at flux of 5 X 10 _° Unprogrammed scan direction reversals Failure to generate planet in view signal Failure to carry out programmed search mode White spots on the picture associated with transients on the video llne Switch to fast exposure speed Increased noise level causing grayish back- ground Decrease in picture resolution False planet acquisition only at high energies and fluxes No anomalous behavior found Damage recovery Star tracking ability regained immediately fol- lowing beam cut off (Front lens element was darkened, degrading transmission by 35% after an inferred total dose of 3000 r.) Immediate recovery follow- ing beam cutoff Recovery achieved after about 20 min for medium flux levels (up to 12 hr for flux levels of 101° ) Recovery following high flux bombardments achieved after about 20 to 30 rain Immediate recovery fol- lowing beam cutoff The Canopus sensor was the only assembly tested capable of giving an analog performance output corre- lated to the electron flux. The other subsystems exhibited a gono-go type of behavior. During irradiation of the sensor, a simulated Canopus star was swept back and forth across the sensor's field of view. The position of 44 JPL TECHNICAL REPORT NO. 32-881 the star was indicated by the potentiometer signal, which was used to drive the X-axis of an X-Y-Y recorder. Inten- sity and roll error voltages were simultaneously plotted vs the roll position of the star. The roll error voltage is a signal proportional to an angular displacement (-1 deg _ 0 < 1 deg) of the star from a plane going through the optical axis of the sensor and roll axis of the spaceea"aft. Disphcements between _ 1 to 3 deg give satn- rated roll error outputs, while displacements > ±3 deg are out of the field of view. A plot of u_'L_sat'aratcd "-v, maximum roll error signal as a function of electron flux is shown in Fig. 39 for an irradiation at 2.25 Mev. The o_ I od E .e. O o G n- X Im 10 II i0 Io i0 9 I0 e 10 7 106 l NEGATIVE SATURATION I.t "_\ VOLTAGE---,,,'/" I FLOSS OF ROLL --[ ATTITUDE C_,NTROL -------5 t RADIATION INCIDENT NORMAL TO THE TOP OF THE TRACKER ELECTRON ENERGY EQUAL TO 2.25 Mev AT THE TRACKER i' ROLL NULL VOLTAGE j [ 1 1 -15 -I0 -5 O 5 ROLL ERROR, v Fig. 39. PerformMnce of the roll error signal under radiation with electronics at 2.25 Mev 10 maximum roll error voltage varies between + 10 and -- 12 v. Loss of roll attitude control occurs when the maxi- mum roll error signal falls below 2 v. Figure 39 shows that this point is reached for 2.25 Mev electrons at an electron flux of 4 X 109 dectrons/cm_-sec. In addition, we note a sizable offset in the star null position voltage at fluxes greater than 109. The polarity of the null offset was a function of the direction of the incident radiation with respect to the sensor. Consequently, null offset vo!_ges were expected to be less of a problem in an omnidirectional flux field. The star intensity signal was also monitored. It was found that when there was no star in the field of view, an apparent star intensity signal was produced whenever electrons were incident on the sensor. Furthermore, it was found that when this apparent star-intensity signal reached six times Canopus brightness, the roll-error signal peak had dropped to the 2-v control loss level. Plots of the apparent star intensity for no star input vs radiation flux rate are shown in Fig. 40 for three electron energies. In this ease, the electrons are incident 32 deg from the normal to the top of the sensor. This direction, which permits the electrons to reach the photoeathede while traversing the minimum absorber thickness, was predicted to be the most sensitive direction to eause a radiation-induced failure. In the final analysis, however, there was no preferred direction of incidence. The ener- gies shown on Fig. 40 are electron energies upstream from the 4-mil eopper scatterer. About % Mev should be subtracted from the quoted energy values to give the electron energy at the sensor surface. The Canopus sensor exhibited the only pemmnent damage degradation of all subsystems tested. The front 10 2 z IO t z tY I0 0 I I O 1.625 Mev-32 deg FROM TOP [] 1.5 Mev 32 (leg FROM TOP <_ t25 Mev 32 deg FROM TOP 6 X CANOPUS IS THE POINT AT WHICH ROLL ATTITUDE CONTROL IS NO LONGER MAINTAINABL_ 107 108 I0 9 i0 I0 i0 II FLUX RATE, electrons/cm 2 sec i012 Fig. 40. Apparent star intensity for no star input vs radiation flux rate 45 JPL TECHNICAL REPORT NO. 32-881 t i lens element was darkened, degrading transmission by 35%. The darkened element was made of flint glass and the integrated dose was estimated to be 3000 rad. Note that the data encoder malfunctioned at electron energies which could not penetrate the flight case, indi- caring that bremsstrahlung produced by the electron deceleration in the scattering foils and the flight case contributes significantly to the system's radiative degra- dation. Such bremsstrahlung produced in the scattering foils acts to make this test more conservative. The planetary scan subsystem was checked for its ability to search for, and lock onto, a simulated planet located outside the vacuum chamber, and to generate a planet-in- view signal. In this case, the planetary scan sensor generated an error signal when the planet was off the sensor axis. The error signal applied to a servo loop resulted in driving an electric motor in the direction re- quired to keep the planet centered in the field of view. As Table 11 shows, failures occurred in all the opera- tional modes monitored. A phenomenon peculiar to the planetary scan and television subsystem was the long recovery time after each failure induced by the radia- tion. This was found to be due to a slowly decaying radiation-induced noise level in the p-n junction sensor. Radiation sensitivity to electrons was energy independent. The television camera was focused onto a test pattern outside the vacuum chamber during irradiation. A de- crease in picture resolution, grayish backgrounds caused by increased noise level in the preamplifier, and large transient pulses on the video signal line were seen. The large pulses caused white spots to appear on the tele- vision picture and were, also, responsible for switching the camera to a fast exposure speed. The threshold failure level was 5 × 109 electrons/cm2-sec, independent of the electron energy. The narrow-angle Mars gate was found to malfunc- tion only in relatively high flux fields at energies above 2.5 Mev. The sensor is a cadmium sulfide photoconduc- tive cell that simply changes resistance with incident light level. A two-transistor trigger circuit comprises the remainder of the circuitry. The failure modes were the appearance of a false planet-in-view signal at the flux levels shown in Fig. 88 or, at lower fluxes, the in- ability of the gate to drop out after the Mars simulator was switched first on, then off. The gyro and attitude control circuitry was the only subsystem that was totally unaffected by the electron irradiation as far as we could detect by overall system performance checks. It is felt that this is mainly due to the fact that all the amplifiers used a very high degree of feedback. Except for the darkening of the Canopus sensor optics, there were no effects that could be attributed to a cumu- lative radiation dose. 4. Summary The extent of the bremsstrahlung produced by the electron beam in the scattering foils and its effect on the electronic subsystems is, at present, an unknown factor. However, the bremsstrahlung does act to make the test conclusions conservative. Rough bremsstrahlung measurements have subsequently been made, under contract," in a somewhat similar irradiation geometry. The dose and flux measurements were also made in a plane 26 in. from a 4-mil copper scattering foil. It was found that for electron energies lower than 1.7 Mev, the dose rate due to foil bremsstrahlung was < 1% of the dose rate due to the electrons, alone. However, the ratio was seen to increase rapidly to a value of 14% at 1.95 Mev, which is the high energy limit of the contracted facility; so the region above 2 Mev was not explored. The accumulated radiation dose given to the items in this test is higher by several orders of magnitude than the accumulated dose intercepted by Mariner IV in traversing the Van Allen belt, the solar wind, and cosmic radiation. The tests indicated that the Mariner IV space- craft should be capable of operating reliably and con- tinuously for at least 1 hr in electron flux fields of 109 electrons/cm_-sec at electron energies between 0.5 and 2.7 Mev. Since this flux was much higher than the practical predictions of Martian trapped radiation fluxes, the project decided not to modify the nominal encounter sequence. The wisdom of this decision was borne out by the charged particle and magnetic field measurements made by Mariner during fly-by. The Mariner measure- ments detected no trapped radiation belts, and the Martian magnetic moment was shown to be < 3 × 10 -4 times that of the Earth's. C. Electromagnetic Interference Test Activities Three tests were performed in the electromagnetic interference (EMI) area during this post-launch period. One of these tests, the RF signature test, had been *At the Statitron facility of Atornics International, Division of North American Aviation, Inc., Canoga Park, California. 46 JPL TECHNICAL REPORT NO. 32-881 waived during the test operations because of limited time available; the results were expected to benefit future spacecraft programs, but were not considered mandatory for a successful Mariner mission. The second of these tests was performed to verify that peak power RF levels expected to be encountered on the launch pad and during the launch phase would not deteriorate the spacecraft pyrotechnics. The third test investigated the _noise that had been observed in space chamber PTM tests when the attitude control gas jets were energized. It had been feared that ff the noise appeared during the encounter sequence, some commands or information would be lost. Two tests that were suggested to the Project Office for further investigation of possible problem areas were not performed. These two tests were: (1) a complete umbilical and separation connector transient test and (2) high- voltage tests that would simulate breakdowns of accumu- lated voltages on the spacecraft/shroud configuration. The tests were not performed because it was felt that the information gained would not be directly applicable to future spacecraft programs. The test of RF noise at the transmitter frequency was not performed because it could not be carried out with- out disturbing the flight configuration. Its performance would not yield valid results ff the flight configuration were changed, and no other test method was devise& The problem was one of making a noise measurement in the presence of the transmitted high-level RF signal. 1. Spectrum Signature Test of PTM No major problems were encountered during the PTM spectrum signa_re test. A frequency range from 80 Hz to 10 GHz was surveyed while the PTM was operated in the cruise mode. The cruise mode was chosen for this test because a majority of the spacecraft functions would be active at this time. The encounter mode, in which the science systems would be in operation, would have been preferred, but that mode could not be kept on for the total test time. Results of this test will be useful in plan- ning test requirements for future spacecraft programs. Measurement was made of the RF environment to which the spacecraft was subjected on the system test complex; this measurement may be valuable in determin- ing Voyager program RF environmental test levels. The RF contribution by the spacecraft was also mea- sured. The exact level generated by the spacecraft could be of importance-for example, in the event an electro- magnetic measurement experiment were included. Such a measurement could be severely restricted by the am- bient RF level; in such a ease, the measurements must be made with the spacecraft in a shielded enclosure that has a low ambient RF level. Since adequate facilities were not available at the time, a complete measurement of the spacecraft RF levels was not made. A spectrum signature measurement, a Mariner program _,a tho v-r_ bet:w.ee.n_ __1 and 30requirement, was m,_ue on ....... April 1965. The findings will be directly applicable to the Voyager program. The results will be used to prepare allowable limits for electromagnetic radiation output from future spacecraft, and the measurement of radiation present on the spacecraft may aid in the design of radiation-measurement experiments to be carried on the spacecraft. A block diagram of the test configuration is shown in Fig. 41. The antennas were positioned 10 ft away because a closer distance would not allow some of the more directive horn antennas to view the complete spacecraft. The spacecraft was first conditioned according to the following list, though not necessarily in the order shown. 1. Spacecraft power on 2. Simulated solar power, Mars mode 8. Data encoder Mode 2 (8½ bits/see) 4. Radio: traveling wave tube, TWT (power ampli- fier A), exciter A, transmitter high-gain, receiver low-gain antenna, RF power up (DC-8) 5. DC-1, cruise on, encounter with scan inhibited, DC-24 N BUILDING WALL .,,,,,,__BAyPTM (MOUNTED ON TEST FIXTURE) S IO fl . ANTENNA ND TRIPOD WALLS _ k,- BUILDING WALL Fig. 41. Block diagram of test configuration 47 ¢JPL TECHNICAL REPORT NO. 32-881 6. CC&S: L-2, L-3, set for attitude control; MT-7 set to turn encounter science on 7. Attitude control: Sun gate, dc power; DC-15 and DC-18 in effect 8. Video storage: 2.4-ke power on (not recording) After the spacecraft was conditioned, it remained in this mode of operation while measurements of the radiated electromagnetic fields were made. The noise and field intensity meters and associated equipment were operated in a normal manner to make the measurements. The measured voltages at the receiver input were later converted to field strength. The greatest problem was identifying the source of radiation. Because of space limitations in the shielded room, it was not possible to perform the test there. As a consequence, the test was performed in a location that has a certain ambient electromagnetic field not contrib- uted by the spacecraft. Signals from local stations and some other sources can be identified by monitoring the received signals with earphones. These signals were not recorded. The remainder, however, may not necessarily be assumed to be due to the spacecraft. When the mea- surement of a frequency band or set of bands on the receiver was completed, the spacecraft was turned off, OSE was left on, and the ambient field was measured. In some cases, the reading had increased by a few db for this measurement. This is typical in a normal un- shielded environment because of variation in man-made and natural radiation, so it caused no concern. However, any spacecraft fields lower than the ambient cannot be measured. Because of various other commitments of the space- craft and test team, the spacecraft configuration was not exactly the same during each day's testing. However, deviations were not considered to be serious. The results are presented in Figs. 42 to 45. With each graph of RF noise measured with the spacecraft on, there is a corresponding graph of RF noise measured with the spacecraft off, but with as much OSE on as possible. In some cases, it can be seen that the spacecraft contributed little or nothing in excess of the ambient noise level. In other cases, it appears that the spacecraft reduced the ambient noise level; in fact, the variation of the ambient could be on the order of ___5 db, which makes interpre- tation of signals near the ambient very difficult. In the frequency region of 1 to 10 GHz, the ambient noise level was low and was not measured. 130 HO 9O 70 5O 30 zL .o 130: "o IIO 90 70 50 3O Io_'SPAC_-CRkF'T'ONDI_OLEFRO'BE I /__\ _ t MEASUREMENTS - 8-Hz BANDWIDTH 60-Hz BANDWIDTH (b) AMBIENT PROBE I I 10 3 104 FREQUENCY, Hz 10 101 102 ;PACECRAFT OFF DIPOLE \ ) Fig. 42. Spectrum signature, dipole probe In the frequency region 50 kHz to 1 MHz and near 10 MHz, the spacecraft apparently emitted broadband RF noise above the ambient noise level. In the frequency region of 50 to 300 MHz, much more broadband noise was recorded with the spacecraft on than with the spacecraft off; more continuous wave (CW) signals were also generated. In the frequency region 1 to 10 GHz, four signals were obviously the first harmonic (funda- mental), second, third, and fourth harmonics of the spacecraft transmitter (2295 GHz). Near these harmonics were apparent sidebands, spaced primarily at intervals of approximately 4-150 MHz on either side of the harmonics. The noise amplitudes shown in the figures are in db above 1/_v/m (CW) or db above 1/zv/Mc/m (broadband) because they are the standard units used in electro- magnetic compatibility (EMC) and in standard EMC specifications. 48 JPL TECHNICAL REPORT NO. 32-881 140 120 I00 8O 6O I00 8o rE _ I I 1BROADBAND NOISE I i I i t II I ! III (o) SPACECRAFT ON 41-in. ROD ANTENNA 1 I Iit I I I l I I I i I I I I -- BROADBAND I NOISE i b I I I 0 10-2 2 4 6 I0 -I 2 4 6 100 2 4 6 I01 2 FREQUENCY, _ MHz Rg. 43. Spectrum signature, rod antenna 1140 120 I00 8o I I J 1so t ) 40 120 ( t E 140 / i 5o! I J J J 4oi 0 [] SPACECRAFT ON VERTICAL POLARIZATION HORIZONTAL POLARIZATION APRIL 22, 1965 ? I I C) ? , : ,r j I ) ( I I ol , 1 i .JlJ 2 II !" 4 FREQUENCY, _ GHz ( IOl I II !!I [11, lill ) ( 1 lJ 6 ( Fig. 45. Spectrum signature test results, spacecraft on vertical and horizontal polarization ) T I <_ 11 I ! I 8 9 I0 31 JPL TECHNICAL REPORT NO. 32-881 ¢ 2. RF Simulation Test With Peak RF Power Densities The purpose of this test was to determine the effects of peak RF simulation levels on the spacecraft pyro- technics. Some of the RF levels had not been used during previous testing because average power levels had been employed. One other test with live pyrotechnics and RF simulation had been performed on the PTM space- craft on 4 and 5 May 1964. That test employed an S-band RF generator that provided a radiation of 100 w from the spacecraft omni-antenna. The shroud was both on and off for that test. No pyrotechnics had been deto- nated or showed any degradation. Reservations had been expressed, though, on the validity of the test because two of the RF sources simulated were simulated with an average CW level instead of the peak level. To increase the confidence on the test safety margin, the test was performed again on 16 April 1965 using peak power density levels. Continued revisions of the RF simulation levels using information obtained after the first PTM test determined that some of the RF levels would now be lower in magni- tude and that two of them, Atlas pulse and rate beacons, would be greater. Decision was made that the higher Arc_AZUSA I r RADAR TRANSMITTER _.J GR-}220 I I LA 24330 I [ TWT AMPLIFIER ALFRED MOD 5-542 LA 34916 25-ft SPIROFLEX AGENA-RADAR TRANSMITTER GR-1220 ALFRED MOD 5-6866 IKS-1271 H TWT AMPLIFIER i25-fl SPIROFLEX LA 33694 COMMAND DESTRUCT TRANSMITTER SIERRA 2130-470 LA 24527 25-ft RG 9 B/U ] DIPOLE A ASGU,DANCE1 [D'RECT'ONALIt TRACKSUBS STEM' O LERj l TRANSMITTER / _HP GRI220, LA 5388[ / -- -- _ x75oo ATLAS GUIDANCE RATE SUBSYSTEM TRANSMITTER GRI220, LA 53881 25-ft SPIROFLEX r_ TWTAMPL'F'ERI ALFRED MOD 527 _ DBG 520 I LA34196 1L-.. _ -- / db 25-ft RG 214/UFATLASTM_I _TTENUATOR IBENDIX JDEL TXV-20 DIPOLE Fig. 46. Simulation equipment level in all cases should be used. Therefore, the test levels used were those used previously, with the exception of the levels for the Atlas rate and pulse beacons. The simu- lation levels, antenna distances, etc., are shown in Table 12, and the simulation equipment setup is shown in Figs. 46 and 47. No pyrotechnic devices were activated by the simula- tion sources during the test. The pyrotechnics were ex- amined and no degradation was observed in either the pinpuller or midcourse motor valve squibs. 3. S-band RF Noise Measurements on PTM in Spacecraft Assembly Facility The objective of this test was to determine whether noise at the spacecraft receiver frequency observed during space chamber tests on the PTM was still present when the attitude control gas valves and other functions were actuated. If noise had been observed, methods were to have been investigated that would determine how to ATLASTRANSPONDERGR_I220AZUSA / ATTENUATOR3db 25-ft FOAMFLEX ATLAS GUIDANCE L [TRACK SUBSYSTEM DIRECTIONAL TRANSPONDER I I I COUPLER PRO 703 H P. ATLAS GUIDANCE L-_--- RATE SUBSYSTEM /TRANSM LTTER ATTENUATOR PRD "/03 - 20 db I TWT AMPLIFIER I MiC_ MODEL 5-542 _ 25-ft FOAMFLEX AGENA TM TRANSMITTER UNITED ELECTRO DYNAMICS MODEL TR-IO AMPLIFIER AMPLIFIER LE TELECHROME BENDIX MODEL 1460M MOOEL TAV-8 I 25-ft RG9 B/U AGENA TRANSPONDER I AVERAGE SOURCE L GR-1220 25-ff RG 214/U POLARAOCA-M_ AND REFLECTOR ]AGENA TRAN SPON DER _/_l PULSER / PEAK I ..... l JPL-BUILT / SOURCE AVION 149A _ I (MODIFIED) SUBCARRIER _ SUMMING OSCILLATORS AMPLIFIER JPL-BUILT TO ATLAS TM I TO AGENA TM Rg. 47. Modulation source for Atlas TM-1 and Agena TM sources 52 JPL TECHNICAL REPORT NO. 32-881 Table12.RFsimulationtestparameters Electromagnetic interference sol/rc8 Agena radar transmitter (FPS-16) Atlas Azusa radar transmitter Atlas guidance track subsystem transmitter Atlas guidance rate subsystem transmitter Command destruct transmitter r-mquency, Gc 5.690 5.060 Classified Classified Classified Atlas telemetry No. 1 Atlas Azusa transponder 0.2299 5.000 Atlas guidance track subsystem transponder Atlas guidance rate subsystem transponder Agena telemetry Agena transponder average source Agens transponder peak source, 500 pps Classified Classified 0.2_3 5.765 5.765 Antenna power, mw Antenna and gain 159.0 Horn" 15 db 2000.0 Horn 10 db 252.0 Horn 15 db 25.2 Horn 15 db 173.0 Half-wave dipole 2.14 db 87.2 Half-wave dipole 2.14 db 35.8 Horn 15 db 200 Horn" 17 db 0.06 Horn" 17 dE 26700.0 YAG1 Y103-N 12 db 6.28 Reflector & Horn feed 24 db 540 W Horn 259 MW 10 db 600 pps 0.8 Fsec Antenna to spacecraft distance Required meters 1.414 1.5 2.9 1.85 Actual measured meters !.4!4 1.Sm 1.5 2.9 1.85 Antenna azimuth Po_trer de.s_y required, mw/m 2 100.0 795.0 640.0 64.0 2.52 1.26 40.0 795.00 0.25 4000 32.0 126,000 Power density (actual) based on calculations using actual power antenna & distance, row/m: 100.0 795.0 640.0 64.0 2.52 1.26 c 40.0 795.0 0.25 40OO 32.0 126,000 •One antenna for two sources. bPe¢ JPL Procedure CM 409.00, p. 14. elt was not possible to set the antenna for this source at the required 3-m distance. A 1.5-m distance was used and a 6-db pod was used to decrease the antenn_ power accordingly. minimize any injurious effects during the encounter. The test was performed in the spacecraft assembly facility (SAF) because of the requirement that the PTM be available for backup tests during the flight of Mariner IV. The frequency translation system of the S-band re- ceiver was used in conjunction with a magnetic tape recorder, and a frequency band of 120 kHz centered about 175 kHz was recorded during the test. The tape recording was subsequently analyzed; apparently some change had been made to the spacecraft, or a change in the test configuration had occurred, since no noise was observed during the test or in the subsequent analysis of the tape recordings. 53 JPL TECHNICAL REPORT NO. 32-881 During the PTM space simulator mission tests, the omni-antenna output was monitored for the presence of noise in the spacecraft receiver frequency to comply with the test and operation plan requirements. Low-loss foamflex coaxial cables and a duplexer were used to connect the omni-antenna output to the EMI receiver console. At various times during the 10-day period of the test, measurements were made and some predetection magnetic tape recordings (in a bandwidth from 25 to 300 kc) were made of the omnidirectional antenna output. It was determined that when the valves were operated, noise bursts were observed with the test re- ceiver. The noise was recorded, and later analysis of one of the longest bursts indicated that the level exceeded the allowable maximum noise level by approximately 20 db during the 50 msec of its occurrence. It was requested that tests with the PTM spacecraft in the space chamber be performed to determine more fully the noise characteristics and effects on the space- craft receiver at encounter. Some objectives of this test were: to determine the mode of noise generation, to determine the magnitude of the noise at the spacecraft receiver input, and to arrive at a method of living with the noise bursts if they occurred during encounter. Certain radio tests during gas let firing were suggested to determine adverse effects. Because the noise bursts were not observed with a horn antenna, even at a close proximity, the radio tests were not performed. Because of the requirement that the PTM spacecraft be maintained in a ready condition for Mariner IV backup exercises and investigations prior to encounter, the PTM was not placed in the space simulator for the noise measurement test. An analysis had shown that a test performed in the SAF would yield a more sensitive noise receiver system sensitivity than the test configura- tion at the space simulator. Therefore, the requirement that the PTM remain in SAF did not adversely affect the noise test-except for the possibility that the simulated space conditions would be a factor in the creation of the noise. a. Receiver system. The S-band noise measurements were made with a sensitive receiver. The noise receiver configuration is shown in Fig. 48. The noise receiver is a double conversion device with stable oven controlled local oscillators. A parametric amplifier with a low-noise figure is used for the spacecraft receiver frequency. Two preselectors are used, one for the center frequency and one for the spacecraft receiver image. The second IF frequency (at 9.56 Mc) is inserted into the receiver. The receiver final IF magnitude at 455 kc is measured with a vacuum tube voltmeter. The system noise temperature for a test configuration in SAF is lower than for a configuration in the space simulator (Appendix B). The system temperature in the SAF configuration is calculated as 945°K, and it is calcu- lated as 81880K for the space simulator. Although the space simulator walls are cooled by liquid nitrogen, the long coaxial cable length with several db of loss causes that configuration to yield a higher system tem- perature. The differences in receiver system temperatures yield a sensitivity margin of 5.3 db, and since it is more difficult to conduct a test in the space simulator, the SAF configuration was preferred. b. EMI noise receiver calibration. During the survey period, the noise receiver was calibrated with the S-band calibrator unit. This unit included a crystal oscillator encased in an oven-controlled stable temperature chassis. The basic crystal frequency was multiplied up to the spacecraft receiver frequency, and the calibrator power output was monitored with a thermistor and a micro- wave power meter. Part of the power was diverted through a directional coupler and attenuated by pre- cision variable and fixed value attenuators. The output calibration RF level was a stable, known value, since the signal path had previously been carefully calibrated. The power supply and power monitor leads were well filtered, and the calibrator case was an integral casting with a tight fitting RF gasketed cover. The variable attenuator dial had been calibrated in position vs attenuation, and a chart listing dial position vs output power was used in the calibration. The calibration technique was to determine the 3-db sensitivity level of the noise receiver and to observe that the frequency read on the noise receiver was correct. Before the tests were started on each day, the gain of the S-band noise receiver was adjusted and the noise figure was measured. This process was repeated halfway through the test and, also, at its conclusion. The receiver remained stable throughout its operation with a noise figure of essentially 3 db. c. Predetection recording. The frequency translator system placed the S-band receiver frequency for the PTM at 175 kHz. The bandwith of interest was 120 kHz cen- tered about 175 kHz. The tape recordings were taken to the Data Analysis Laboratory for analysis. The analysis immediately available was one using a wave analyzer 54 JPL TECHNICAL REPORT NO. 32-881 t CW CRYSTAL CONT ROLL ED ALIB ATOR I POWER ISUPPLY POWER ISUPPLY NOISE RECEIVER SPACECRAFT RECEIVER FREQUENCY I_1_ I I I SPACECRAFTRECE,VERI IIMAGE FREQUENCY I I-I I I I i PRESELECTORII '° II PRESELECTORt--_IMAGE PARAMETRIC IAMPLIFIER AMPLIFIER B J I POWER SUPPLIES I NOMINAL 47.5 Mc _I AMPLIFIER CRYSTAL IOSCILLATOR 9,56 Mc IGNAt IVIDEI t I COLLINSRECEIVER 455 kc ___IF_ 9.56 Mc t FREQUENCYTRANSLATOR I 9,56 Mc r ...... I ICOUNTERLrSTABLER _RENCEi I- I FREQUENCY TAPE RECORDERAMPEX I _I MONITORI -LOSCILLOSCOPEl L .... 300 cps-300 kc Fig. 48. Noise receiver configuration for S-band noise measurements 1 I I I I '1I I I __J 55 JPL TECHNICAL REPORT NO. 32-881 that required a tape slowdown by a factor of eight. With a 20-Hz bandwidth filter, a resolution of 160 Hz at the original frequency was possible. There existed a limita- tion with the wave analyzer; the cutoff frequency was 25 kHz and this would correspond to an original fre- quency of 200 kHz. The data were then analyzed in a bandwidth of 175 (+25, -60) kHz. Since this data reduc- tion was to be merely a preliminary analysis, the cutoff would be acceptable. It was hoped that an analysis could later be made of the entire band. The magnetic tape recordings made during a series of valve actuations, Sun gate initiate commands, etc., were analyzed and calibrations were placed on the tape re- cordings; these included discrete S-band CW calibrator RF levels, an ambient 50 f_ termination, and a gas tube noise source. d. Measurements in SAF. The measurements were performed using the spacecraft omnidirectional antenna and a portable S-band horn antenna. For tests with each antenna, the following attitude control functions were actuated: 1. Yaw and pitch valve firings clockwise (cw) and counterclockwise (ccw) 2. Roll ccw 3. Initiate Sun gate 4. Canopus sensor on No noise was observed when the spacecraft omnidirec- tional antenna was used (this would be the noise intro- duced into the spacecraft receiver). After a period of testing, a horn antenna (with 15.9-db gain at 2113 Mc) was set up on a tripod approximately 4 ft from the space- craft. Again, the several functions were actuated. Only one possible noise burst was observed, but it could not be repeated. Since it was not repeatable, its origin was attributed to some other cause not found. A light-weight horn antenna (with similar gain) was then used to probe about the spacecraft as individual elements were ener- gized: the several valves, the Sun sensor, the Canopus sensor, and Case VII. No noise bursts were seen. The horn antenna, connecting cable, and receiver were veri- fied to be working when the background noise level varied with antenna orientation. Although no noise bursts were observed visually on an oscilloscope at the EMI S-band receiver output, a tape recording was made during the series of valve actuations, Sun gate initiate commands, etc. The results of the analysis for six separate spacecraft events were available a few days before encounter, and an examination of the results seemed to indicate no apparent discrete noise components. The analysis of the magnetic tape was performed at discrete occurrences of the noise test, but it was difficult to determine the exact time of each spacecraft occurrence. This experience indicates that ff future tests of this type are performed, instrumentation will be required that will place an event marker on one of the tape recorder data channels at the time of its execution. e. Conclusions. Since no noise bursts were observed in these tests and some had been observed with the PTM in the space simulator, the following reasons may account for it. 1. The vacuum environment is required for the noise to be generated. 2. Any noise, however small, when in the space simu- lator, is reflected efficiently into the omnidirectional antenna. This is not the condition for a spacecraft in free space, of course. 3. Some change in spacecraft hardware or configura- tion between the two tests eliminated the noise generation. D. Magnetic Stability and Demagnetization Tests While the Mariner solar panels were successfully de- magnetized for the flight spacecraft, their magnetic stability remained an uncertainty. Because of this lack of information, limited magnetic tests were scheduled subsequent to the Mariner IV midcourse maneuver on TA solar panels and on the complete PTM bus. In the first tests conducted in February and March 1965, two TA solar panels were subiected to vibration tests and storage for periods of four weeks in controlled environments in both the magnetized and demagnetized conditions. The tests on the PTM bus were conducted during a week in May followed by a second week in August 1965. In the earlier tests, attempts were made to demagnetize the spacecraft in the presence of the Earth's field, while in the latter tests, the ambient field was reduced by more than an order of magnitude using a special coil system. 1. Solar Panel Magnetic Stability Tests The solar panel stability study, which had as its primary purpose a determination of the stability of the 56 JPL TECHNICAL REPORT NO. 32-88! flight solar panels during Munch operations and inter- planetary cruise, was conducted in four phases: 1. Determination of the effects of vibration in the presence of an ambient magnetic field 2. Storage of a demagnetized panel in the Earth's field 3. Storage of a demagnetized panel in a low-field envirom_ent 4. Storage of a magnetized panel in a low-field en- vironment The vibration tests were conducted to indicate what the effects of launch might be on the solar panels due to vibration in the Earth's field as enhanced by the presence of the launch vehicle and umbilical tower. Unfortunately, the degree of similitude is not good. The vibration levels in these tests are higher, and the ambient field is be- lieved to be more than an order-of-magnitude higher and very- divergent. Vibration fields as high as 10 gauss were measured near the solar panel when it was mounted on the vibration exciter. The field of the launch vehicle and umbilical tower is not so well known. Limited measurements indicated that the magnetic field at launch is only slightly more than the Earth's field, alone. The vibration tests disclosed that magnetization of the solar panels was entirely independent of the frequency of vibration over the range of testing (40 to 2000 cps). A magnetized panel appeared less stable than a demagne- tized panel when shaken on the vibration exciters. The residual magnetization acquired by the demagnetized panels was consistently the same, whereas the magne- tized panel, when shaken, ended with a residual mag- netization that appeared to bear no relation to either the initial magnetization or the shaker field. It appears that the resultant, residual field of a demagnetized panel following vibration is entirely a function of the ambient field from the vibration exciter and the orientation of the panel in the ambient field. The solar panels, with their Kovar strips and wires principally parallel to the long axis of the panel, became magnetized to a considerably higher value with the field in this axis. It was found that the solar panel residual magnetization followed a typical magnetization curve, with satttration occurring at about 20 gauss (which is the expected value for Kovar). In the region of the Earth's field magnitude, the field at the magnetometer would be about 2 _. The storage tests indicated that a demagnetized solar panel is relatively stable when stored either in a low (< 100 _) field or in the Earth's field. A magnetized (penned) solar panel, on the other hand, definitely loses its magnetization with time when stored in a low-field environment. Earlier measurements made on solar panels transported between Pasadena and Cape Kennedy also showed negligible change in field over a period of several months. For these low-field tests, the panel was stored both in a magnetic shield room and in a single-axis spherical coil system; because of the size of the solar panels, the ends of the panels were exposed to somewhat higher fields; however, changes ha thc residual magne- tization of the solar panels was negligible in all cases except the stored magnetized panel. From the above study, it can be concluded that the magnetic field due to a demagnetized solar panel should not change during interplanetary cruise, but could change a limited amount on launch. It is estimated that the change of field of each solar panel on launch would be less than 2 _, even under worst-case conditions. 2. Proof Test Model Bus Demagnetization Study Time did not permit a demagnetization of the Mariner HI or Mariner 1V bus structures before launch. With the successful demagnetization of the solar panels, there was renewed interest in the possibilities of de- magnetizing the remainder of the spacecraft; so a study was scheduled on the PTM spacecraft. A 1/20 cps alter- nating, decreasing field was used for the spacecraft demagnetizing to avoid inducing damaging voltages in the spacecraft circuitry. The first week of the study consisted of attempts to demagnetize the spacecraft in the Earth's field. Demag- netization was attempted in successively increasing initial field steps of 5, 10, 20, 40, and 80 gauss in each of three orthogonal axes. The Z axis component of the spacecraft field, which is the predominant one, was decreased in steps from 301/_ down to 17 _. At each initial field level, the spacecraft was demagnetized in two perpendicular axes in the X-Y plane, followed by mapping, and in the Z axis with the order alternated at successive levels. Even though the initial field was doubled at successive levels, the change in remanent field was less in the instances where successive demagnetizations were in the same axes, than when the initial field was the same and the orienta- tion of the demagnetizing field changed. This indicates that it will probably be necessary to demagnetize a spacecraft in more than one axis, except where the de- magnetizing field can be aligned with the residual mag- netization of the spacecraft. _7 J P L TECHNICAL REPORT NO. 32-881 Following this partial demagnetization, it was con- cluded that the saturation level of some spacecraft material, such as the hard magnets, had not been reached and that, in order to further demagnetize the spacecraft, it would be necessary to reduce the ambient Earth's field. This latter course of action was followed Fig. 49. Mariner X-Y plane demagnetization facility with Earth's field bucking Helmholtz coils in August 1965. Available 10-ft-diam Helmholtz coils were adapted to encompass the spacecraft demagnetizing fixture. For demagnetizing in the X-Y plane, the demag- netizing coil axis was oriented in the magnetic north- south direction and two pairs of Helmholtz coils were arranged to buck out the vertical and north-south corn- ponents of the Earth's field (Fig. 49). With this arrange- ment, the ambient field over the volume to be occupied by the spacecraft hiis structure wzs redgcec! to < 4090 -1, or almost an order of magnitude. In demagnetizing in the Z axis, a single pair of Helmholtz coils were inclined so as to oppose the total Earth's field (Fig. 50). Over the large volume of the spacecraft, this arrangement reduced the ambient field to < 3000 y. In these latter tests, demagnetization was only attempted at initial field values of 40 and 80 gauss. The improvement in field reduction at 80 gauss was very slight. For optimum demagnetization, it was necessary to demagnetize in three axes. By this means, the PTM spacecraft bus was demagnetized so that the remanent field at the magnetometer location was reduced to H , = -4, H , = 0, H , = 4% 7. In the course of these tests, the spacecraft was magnetized by exposure to fields of 25 and 50 gauss. The spacecraft acquired a much higher residual magnetization when magnetized in the Z-axis direction. Following this magnetization, the re- sidual magnetization could be reduced to its minimum Fig. 50. Demagnetization of PTM spacecraft in z axis with Earth's field bucking Helmholtz coils 58 JPL TECHNICAL REPORT NO. 32-881 value by a single demagnetization in the same direction as the earlier magnetizing field. Table 18 smnmarizes the results obtained with the Mariner PTM spacecraft. In the magnetized case, the sign is dependent on the direc- tion of the magnetizing field. Table 13. Mariner C PTM residual field at Magnetometer Condition of PTM spacecraft Before demagnetization +2½ After 40F demagnetization in Earth's field --4½ After 801" demagnetization in Earth's field --4_6, After 4017 demagnetization in low ambient --5_ After 80F demagnetization in low ambient --4 After 251_ magnetization X-Y plane --41 After 50I' magnetization X-Y plane .1.72 After 25I" magnetization Z axis "1"2 After 50F magnetization Z axis --10 "t-3½ ,+301h 0 ,1,19 --1/2 .1.17 0 "1"4 0 -I- 4_ --32 "1"12 --68 "1"29 •1"27 .1.132 --51 --205 E. Proof Test Model, Usage and Disposition 1. From I to 8 December 1964 PTM spacecraft operations during the period 1 through 8 December 1964 were devoted to support of Mariner IV mideourse maneuver preparations and problem analysis resulting from the loss of Canopus acquisition, when the DC-27 command was transmitted at the start of the first midcourse maneuver attempt. Specifically, the tests were: 1. An SFOF mideourse maneuver practice test 2. An SFOF midcourse maneuver using command tapes prepared for the proposed maneuver to verify accuracy of the tapes 3. A series of tests designed to aid analysis of the Canopus loss that occurred at the start of the first midcourse maneuver attempt 4. An investigation of Canopus sensor susceptibility to power system transients a. SFOF midcourse maneuver tests. A simulated mid- course maneuver test was conducted on 1 December 1964 to provide practice and procedure verifcafion for the operations team. Spacecraft operating time was 2.9 hr. As a simulated problem, a shorted squib was inserted in each half of the pyro harness. Spacecraft operations were completely normal. The second SFOF midcourse maneuver test was con- ducted on 3 December 1964 to provide practice for the operations team and to verify the accuracy of command tapes prepared for the MC-3 maneuver. Spacecraft operating time was 2.7 hr. The midcourse maneuver was performed by the PTM spacecraft with commands supplied by the SFOF mid- course tapes. All maneuver and burn durations and polarities were verified by measuring the responses of the PTM. No simulated problems were introduced and spacecraft operation was normal. b. Maneuver abort problem investigation. An investi- gation of maneuver command sequence responses was conducted on 4 December 1964. Spacecraft operating time was 6.5 hr. Objectives of the investigation were to examine pos- sible explanations for the loss of Canopus acquisition at the instant that the midcourse maneuver counter was started by a DC-27 command and to develop procedures for circumventing another maneuver abort ff acquisition loss was found to be inevitable. Three maneuver sequences were completed with Canopus intensity adjusted to the lower limit. Three at- tempts failed to cause loss of acquisition. But it was likely that sufficient torque was produced to cause loss of Canopus acquisition if the flight spacecraft roll were at a limit of the roll cycle. The sequence was repeated with positive and negative roll errors introduced and again with occasional gas valve operation. Canopus acquisition was not disturbed on the PTM spacecraft. To establish corrective procedures to recover acquisi- tion, if lost at the time of the DC-27 command, six varia- tions of the proposed roll command sequence were tried to determine the effect on the spacecraft. Repetitive roll increment commands were issued at short intervals with- out inducing unusual response from the spacecraft. Acquisition was accomplished immediately after roll increment commands. All of the procedures attempted were successfully completed with no apparent spacecraft problems induced in the PTM. 5g JPL TECHNICAL REPORT NO. 32-881 c. Canopus sensor noise susceptibility test. A Canopus sensor noise susceptibility test was conducted on 8 December 1964. Spacecraft operating time was 1.1 hr. The objective of this test was to determine the sensitivity of the Canopus sensor to power system transients gener- ated by attitude control gas valve operation. This test was conducted in support of analysis of Canopus acqui- sition loss during cruise conditions on 7 December 1964. Transient conditions were induced with the Canopus sensor adjusted for various roll error and intensity adjustments. Results of the tests were entirely negative. Canopus acquisition was not disturbed by power noise transients produced by any combination of gas valve operation or gyro on commands. 2. From 9 December 1964 to 11 February 1965 The period from 9 December 1964 through 11 February 1965 was primarily devoted to preparations for the early encounter exercise performed by Mariner IV on 10 February 1965. A series of tests was conducted to investi- gate anomalies observed in the functioning of Mariner IV and to support SFOF early encounter investigations prior to the actual performance (see Fig. 51). The test sequence follows: 1. A science investigation of the Mariner IV Plasma Probe failure mechanism 2. A television encounter 3. An attitude control investigation of gyro turnoff delay reduction, as a function of Canopus acquisi- tion loss interval 4. A science cosmic dust investigation 5. A radio investigation of suspected Mariner IV carrier suppression change, caused by the flight data rate switch from 331/3 bits/sec to 81/3 bits/sec 6. A system verification test 7. A science subsystem to OSE computer and 83.3 kc interface verification 8. A computer data system data reduction demonstra- tion for Surveyor personnel 9. A series of three SFOF science command sequence verification tests 10. Five SFOF scan platform positioning tests I a. Plasma Probe noise test. To determine the effects of an arc in the high voltage circuit, a science Plasma Probe noise test was performed on 11 December 1964. The purpose of this test was to determine if the sus- pected failure of the high voltage bleeder resistor in the Mariner IV Plasma Probe could adversely affect en- counter experiments. High voltage arc simulation was achieved by connect- ing a spark gap across the Plasma Probe cup, to ensure arcing when the voltage stepped to the highest levels. b. Television encounter. A television encounter was performed and encounter data were examined for evi- dence of interference due to arcing. The noise bursts caused the data automation system to issue an erroneous motor stop-start command to video storage and also pro- duced interference in the form of a television picture line displacement. Data automation system synchronization was disrupted occasionally and the data encoder skipped a 200-deck position when Data Mode 2 operation was attempted. Spacecraft operating time was 3.4 hr. c. Attitude control Attitude control examined the gyro turnoff time delay following Canopus acquisition as a function of the duration of the loss of acquisition. The purpose of this test was to improve the explanation of results noted on Mariner IV flight data following brief losses of Canopus acquisition. Results of the test demonstrated that a brief loss of acquisition will result in a turnoff time delay of less than 3-min duration and that a very short acquisition loss interval did not turn on gyros. d. Cosmic Dust Detector. Science conducted a Cosmic Dust Detector calibration test during 5-6 January 1965. Mechanical preparations were completed during the period 28 December 1964 to 4 January 1965 to put the spacecraft in complete flight configuration, except that dummy solar panels were installed. The test consisted of dropping shot on the spacecraft bus, Case IV, and on the dummy solar panels, while observing the output of the Cosmic Dust Detector. Ob- jective of the test was to verify sensitivity and calibration of the Cosmic Dust Detector as related to particles striking spacecraft surfaces. During the test the Cosmic Dust Detector was operated from an external source; no other spacecraft systems were turned on. 6O DATE SPACECRAFT OPERATIONS HOURS OF SIC OPE RA TIONS OPERA TIO NS MARINER PTM TEST AND OPERATIONS SUMMARY LAUNCH & FLIGHT u SUPPORT OF MC-2 PREP OF S/C FOR LAUNCH SUPPORT I ! I I 1080 1090 I00 II II1IZ/64 111/3/64 I 111/_/64 I I I 11/6/64 II I 11/,/64 _¢_1_1¢,4 I I I 1119/64 ...... _{I i ! 1o/26/64 I 11110164 [ 11112164 I[ II _'OLLOW -UI _.- MC-2 PROBLEMS I i I .1110 1120 1130 I : II I I I RATION TESTS BLDG. 144) ill I MECHANICAL PREP FOR VIBRATION II/17/64 ,11118/64 I I I I |11/1316_ I111114/64 III 1,,,_/64i Ill I I SPECIAL TEST PREP TO MC-3 LAUNCH I ti 140 1150 I! i I li 1160 li [ ! .t i [ If70 1180 I I ! ! I NEW GCDLOCKUP PROCEDU,_.Z I SIC COMMAND THRESHOLDS I SFOF-S/C FAILURE MODE EXERCISE II SHIP TA SOLAR PANELS l FLT. SPARE GAS SYSTEIV DE BIT ERROR RATES ON SIDE BAND LOCK SQUIB SHORT PYRO EVENT AT L-! S/C BEHAVIOR AT LOW VOLTAGE SYSTEM VERIFICATION TEST CHECKOUT OSE FUNCTIONS 11119164 [ 11/20/64 11/23/64 II124164 1211164 1213164 1214164 ! 190 1200 [ 1113165 I 1/15/65 II 1/18/65 II ,11191. II h1/20/. II Ill1/,2/65 II III l/26/65 1128165 IZ/11/)4 1129 12/14/64 • 1111165 , II ' iliH : LAUNCH & FLIGH'I ',IIllrilllI fl ,tll III 11 ill I I I 1210 1220 1230 1240 1250 il I I_ II I i _ I I il i i I I I L 1 -SCI TV/PDP-4 IN NEW TWT SPC & A( CONFIRM CORRECT RF ] •CONFIRM SHORT GYRO-O]_ EFFECTS OF PLASMA PRC EFFECTS OF PLASMA PROI CANOPUS INVESTIGATION ANI SYSTEM VERIFICATION TESTS TIM INDICATION OF RF DOWN SIC COMMAND THRESHOLDS FOLLOW UPON TRD P/FR 22 HR SFOF-S/C REAL TIME EXERCISE _5 _./1/6s z/s/65 SUPPORT OF MC-3 [260 1270 _J Reduced RF power in auxiliary oscillator mode Magnetometer response to gas valves-panels up Immediate NAA after DC-25 Qualify GCD-DSIF amplifier modification Follow-up on PTM radio false lock * Non-standard charger mode Receiver acquisition characteristics Maneuver checkout and support Early encounter system verification Early encounter feasibility, procedures checkout, and SFOF exercise. FERFACE CHECKOUT : DRIVES POWER AT 8 BIT CHANGE 'F DELAY PBE ARCING _E ARCING CORRECTIVE PROCEDURE CHECKOUT :_PROBLEM DISCOVERED 28 OCTOBER DURING SYS. VERIF. TEST FOR VIBRATION, RESOLUTION: LEAVE RANGING OFF. SAF P/FR NO. 299 2/lZ/65 2126165 312165 { { IZ70 1280 1290 I I I 3111/65 3/18/6_ 13122 { { IL 1300 1310 32( I III "--2 I MARINER IW FLIGHT SUPPORTCY-I EFFECT ON RADIO AUX. OSC FREQUENCY (l 7.4 HRS} PHOTOGRAPH { SIGNALS FROM { ?Wl and 9W9 { RING HARNESSES. [ PREPARATION FOR PRINTED { CIRCUIT RING l HARNESS { EWALUATION | (13. I HRS} { I. Printed Circuit Interface Test 2. System Verif. Test CM301.01 3. RWXr-FC Comp. Test {3. 1 hrs} MARINER IV FLIGHT S RW_- F/C COMPATIBILITY DEMONSTRATION SYSTEM T 65 31Z4165 3126165 3130165 416165 • I_ ILr. "gl i I v_ 4/9/65 4112165 1330 1340 1350 4113/65 I 411SI65 4/23/65 I 5110165 II 5/24/65 II 5/27/65 II 512816, II 6/1/65 6/Z/65 -1 1400 1410 1420 614165 6114165 6115165 6123 1430 1440 1450 1460 1470 1480 ! Special Tests ._ I '- . _. _,._,g,a_u,e . I-tl' MAG-MAP AND DEGAUSS (NO OPER) i " [3. Raaio J. A/_ _omma_o I S3e.... _°eR'esAD/o_ns_e°mma'n° I _S/C-LAUNCH MODE CONTROL SET COMPATIBILITY _ ........ _r_'// _ TEST (0.7 HRS} Ill I I I I I I I I.SFOF ENCOUNTER TEST PROCEDURE CM503.00 [ I I Z. RFI MEASUREMENTS (SIMULTANEOUSLY)(8.3 HRS) I I ' LIVE PYRO WITH PEAK R.F. SIMULATION (1.4 HRS) I 'sror DC-18 ENCOUNTER TEST (8.3 HRS| I I I SFOF DC-18 ENCOUNTER TEST WITH NOISY CANOPUS I SF2F ..... TRACKER,6.1 RS III I t CANOPUS TRACKER CALIBRATION ° PREP FOR SFOF MODE TEST TEST (I.0 HR) III [ MARINER IV FLIGI-IT SUPPORT RADIO PATH LOSS TEST 9.9 HRS) ill l RADK) SPECIAL TEST. HIGH RESOLUTION SPE PLOT (7.5 HRS) RWV-FIC COMPATIBILITY TEST {0. Z hrs) PRINTED CIRCUIT RING HARNESS PHOTOS COMPLETED SIN 3 (7.4 hrs) PRINTED CIRCUIT RING HARNESS PHOTO EVALUATION CO MPLETED SIN (7.3 hrs) 1. DATA ENCODER SPECIAL TEST Z. RADIO SPECIAL TEST {0.5 hrs) PRINTED CIRCUIT RING HARNESS PHOTO EVALUATION S/N 1 {9. 8 hrs) PPORT CANOPUS UPDATE EFFECT ON D/E (4.4 HRSI (1.5 HRS} _-ST PTMCM 312.00 (Z. 3 HRS} (9.2 HRS) SPECIAL TESTS 1. Radio Uplink Lock Loss Prob. Investigation I2. Video Storage Investigation 3. Command Lockup 4.Science Cosmic[ l- Dust Detector I / Inves.(24.5 hrs)[ I SFOF ENCOUNTER | TEST | (16. 5 HRS) [ SPECIAL TESTS I. RADIO 2. SCIENCE l, SFOF SCAN POSITION TEST (6.3 HRS) SCIENCE SPECIAL TEST TRAPPED RADIA DETECTOR CALIBRATION BY EXPERIMEI &/-3 165 6124165 I I I 490 ! 500 I 61Z8165 6130165 711165 71Z165 I 1510 II I _/6/65 718165 C_-MND SPECIAL [TEST I Inves. Cmnd. ---_ Lockup Ability ] as a Function of I Radio SPE Off- Lset(1.7hrs_ F---I I / 520 1530 |-_FOF FINAL 1540I II OPERATIONAL L! READINESS TEST "I Simulated Prob- L_lems (13. l hrs) POWER SPECIAL TESTS Inves, Effects of Reduced So- lar Power Limits on Encntr Power Modes (3.8 hrs) SFOF SCAN POS.TEST 6.7 hrs) ' -SPECIAL TESTS I _axiio Verification 2. Power Inves. Effects of Various Encounter Load Combinations 3.Scan Inhibit Logic Cmnd m Nominal Internal Test (3.3 hrs) SFOF ENCOUNTER BACKUP MODE TEST (15.6 hrs) SFOF ENCOUNTER BACKUP MODE TEST (9.0 hrs) PECIAL TESTS |.Radio Uplink Lock Loss Problem Inve stigation 2. Science Scan Inhibit Switch Inves. $.Power Inves. Effects of Various Encounter Load Combinations and Results of Load Sharing on S/C Subsystems {25.2 hrs) L6 HRS} TION ITER {3.9 HRS) 7/9/65 7/IZ/65 550 1560 1. 7114165 8/3 !i"7/20165 to 7/29/65 8111165 to 8/IZ/_ 18113/65III _1!6/65 to 8/20/_5 i, ii I 7130/65 to 8/9/65 1570 1580 1590 1600 1610 LOAD SWITCHING 1 TSANSIENTS,, 1 Print_ 2. COMMAND I [[ 2. Radiol i INVESTIGATED i LOCKUP I RADIO IN] CAPABILITY NOISE INV_ DURING RADIO SPE CHANGE RADIO SPECIAL (4.0 HRS) SYSTEM VERIFIi SFOF MAGNETIC MAPF FINAL Degauss, Rem_ OPERATIONAL READINESS SPECIAL TESTS ON P! TEST UNDER 1. Found PTM wiring NOMINAL Z. Investigated noise b, CONDITIONS 3. Investigated noise di (13.1 fiRS) 4. Investigated spaceci JPL TECHNICAL REPORT NO. 32-881 !/6s Si6S /1°/65 i 9/lsl6s J 9/Z3165 II II 1630 I I 1640 L TEST TERMINATED AND SPACECRAFT PLACED IN STORAGE RADIO ENCOUNTER FREQUENCY SHIFT INVESTIGATION (5. Z HRS) FINAL SYSTEMS TEST PROCEDURE CM 300.03 (13.3 HRS) I "IDEO STORAGE - DAS INTERFACE NOISE TEST (3.6 HRS) INTEID I CIRCUIT RING HARNESS EVALUATION (l. 3 HRS)I l _E INVESTIGATION (4.8 HRS) I I _d circuit ring harness evaluation. uplink carrier level suppression tests (3.4 hrs) 1 1 FESTIGATION CAVITY AMP. (4.0 HRS} :STIGATION ON SPACECRAFT SYSTEM RETURN LINES (7, Z HRS) TEST Anamoly /_nvestigation (Z. 8 hrs) .EATION TEST - CM 301.01 (7.7 HRS) _ING g, Remap )WER SYSTEM NOISE rror in ground return system. tween spacecraft and power ground. e to OSE-spacecraft interconnection. aft system noise due to facilities power. Fig. 51. Mariner PTM test and operations summary JPL TECHNICAL REPORT NO. 32-881 e. Radio carrier suppression test. A special radio test was conducted on 8 January 1965 to determine ff the change of data encoder data rate from 33½ bits/see to 8% bits/sec could be detected by a ground station as a change in radio carrier power level suppression. The reason for this test was to confirm or refute reports of a change in carrier level suppression on the Mariner IV signal coincident with the MT-6 event. As a result of these tests, the apparent change of Mariner IV signal level was attributed to the L- to S-band converters at Woomera and Johannesburg. Spacecraft operating time was 1.6 hr. f. System verification test. A system verification test was conducted during the period 13 January to 19 January 1965. Spacecraft operating time was 13.9 hr. The objec- tive of the verification test was assurance that all space- craft subsystems were in flight condition, following the mechanical rework necessitated by the cosmic dust test, and as preparation for the early encounter exercises anticipated for support of the Mariner IV flight sequence. Although several significant events occurred during the test, none were caused by spacecraft malfunctions. An external power system surge resulted from an error in shutdown procedure and caused the radio TWT am- plifier power supply to fail. Procedure changes and operator instruction were instituted to preclude power transfer surges during future operations. During the playback portion of the test, video storage measured the noise generated by the foil section at end-of-tape to determine ff the PTM tape had deterio- rated from excessive operation. No detrimental noise was present. Following installation and checkout of a new TWT amplifier and power supply, the verification test was completed without incident. g. Science encounter test. A science encounter test was conducted on 20 January 1965. Spacecraft operating time was 2 hr. Objective of this test was to verify the 83.3-kc interface between the television subsystem and the PDP 4 recording equipment in preparation for the early encounter exercises. h. Spaceera[t-computer data system interlace der_ onstration. A spacecraft-computer data system interface demonstration test was conducted on 22 January 1965. Spacecraft operating time was 1 hr. The purpose of this test was to demonstrate the application of the com- puter data system (CDS) to spacecraft telemetry and direct access measurement evaluation, to enable the Surveyor personnel to determine the usefulness of CDS on their program. i. SFOF eady encounter command sequence verifica- tion tests. Three early encounter command sequence tests were conducted from 27 to 29 January 1965. Space- craft operating time for all three tests was 14.6 hr. "Flae objective of these tests was to enable SFOF personnel to test every sequence of commands that might be em- ployed during the early" encounter exercise and to determine if any combination commands could produce a deleterious effect on the spacecraft. All command com- binations were found to be acceptable and usable without restriction. i. Space Flight Operations Facility tests. Five scan platform positioning tests were conducted during January and February 1965. Total spacecraft operating time was 30.2 hr. 3. From 12 February to 14 July 1965 The pre-encounter phase of PTM operations was pri- marily devoted to support of Mariner IV flight encounter preparations. In addition, several series of tests were conducted to analyze nonflight problems and to evaluate printed circuit ring harnesses substituted for the 9Wl and 9W9 ring harnesses. Flight support tests consisted of twelve SFOF practice encounter tests and fifteen flight simulation tests that were conducted to verify or analyze Mariner IV data. Twelve special tests were completed to investigate spacecraft problems that were not directly associated with Mariner IV flight conditions but that needed solu- tious for satisfactory spacecraft flight support and future design verification. The test sequence for this period is shown in Table 14. a. Demonstration system test. A demonstration system test was performed on 12 February 1965. Spacecraft operating time for this test was 2.3 hr. The objective of the test was to demonstrate Mariner system test facilities and computer data reduction methods to the Surveyor program personnel to assist them to adapt the Mariner techniques to their needs. Spacecraft operation was nor- mal and the demonstration was concluded satisfactorily. 63 JPL TECHNICAL REPORT NO. 32-881 No. 1 2 3 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 Date 2/I/65 2/26/65 3/2/65 3/4/65 318165 3/9/65 3/10/65 3/11/65 3/15/65 3/22/65 3/26/65 4/7165 4/9/65 4/12/65 4/13/65 4/15/65 4/16/65 4/20/65 4/21/65 4/22/65 4/23/65 5/10/65 5/12/65 5/27/65 Table 14. Test sequence, 12 February to 14 July 1965 Test Demonstration system test Procedure PTM CM 312.00 Ground command/flight command compatibility Canopus tracker update/data encoder deck skip investigation Investigation of downlink signal dropout on flight spacecraft coincident with CY-1 command Canopus sensor update/data encoder deck skip investigation continued Investigation of effect of CY-1 on transfer commend voltage; continuation of test No. 4 Continuation of test No. 4 Continuation of test No. 4 Radio calibration of auxiliary oscillator frequency vs transfer command voltage Ring harness interface test; investigation start for printed circuit ring harness evaluation, continuing through 4/6/65 System verification test following installation of printed circuit ring harness Radio plot; no signal static phase error (8PE) Radio SPE fine calibration Radio investigated path loss through open circulator switch, pert of interferometer effect investigation Modified Canopus sensor calibration; preparation for SFOF test SFOF DC-18 encounter test SFOF DC-18 encounter test Live pyro-peak RF simulation SFOF test and RFI measurements RFI, RF Signature, attitude control command sequence investigation, radio investigation of PTM uplink lock loss anomaly RF signature RF signature Launch mode control set compatibility Magnetic mapping and degauss investigation, continuing through 5/26165 SFOF encounter test No. 25 26 27 28 29 30 31 32 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 49 50 Date Test 5/28/65 6/1165 6/2/65 6/3/65 6/7/65 6/8/65 6/9/65 6/10/65 6111165 6/14/65 6/15/65 6/17/65 6118/65 6/21/65 6/23/65 6/24/65 6/25/65 6/28165 6/30165 7/I/65 7/2/65 7/6/65 7/8/65 7/9/65 7/12/65 7/14/65 Science Trapped Radiation Detector test SFOF scan position test Radio uplink lock anomaly investigation, science television voltage investigation SFOF encounter test, continuing through 6/4/65 Radio uplink lock anomaly investigation Radio uplink lock anomaly investigation, science DAS investigation Flight command Iockup investigation Radio uplink lock anomaly investigation Video storage count-two circuit transient response investigation Science Cosmic Dust Detector noise investigation, flight command Iockup investigation SFOF encounter backup test Radio uplink lock anomaly investigation, scan platform inhibit investigation Radio uplink lock anomaly investigation Power investigation of low primary power voltage operation, continuing through 6/22/65 Power investigation of encounter load share conditions SFOF backup mode test Calibration of special instrumentation installed for SFOF tests SFOF backup mode test Radio subsystem verification, power investigation of encounter load share conditions, scan platform inhibit logic investigation SFOF scan position test Power modes test for encounter loads SFOF operational readiness test Flight command Iockup capability investigation SFOF final operational readiness test Power transient investigation, flight command Iockup investigation Flight encounter simulation 64 JPL TECHNICAL REPORTNO. 32-881 b. Madner IV flight support tests. A series of tests in direct support of the Mariner /V flight data analysis effort was conducted during this phase of test operations. Spacecraft operating time for the series was 74.3 hr. The first test was an attempt, on 2 March 1965, to duplicate telemetry indications of a failed suppressor diode on a relay operated by the MT-1 command from CC&S. Although a data encoder deck skip apparently occurred at the flight spacecraft when the Canopus sensor cone angle was updated, no similar effect could be produced on the PTM. Commands to the PTM were issued with the relay driver circuit in normal configura- tion, as well as with the suppressor diode disconnected, but no deleterious effects resulted. Five radio tests were completed between 4 and 11 March 1965 to investigate a frequency shift of the flight transmitter signal coincident with the issuance of CY-1 commands. The results of these tests proved that the radio transfer coinmand voltage experiences a level change concurrent with the first CY-1 command issued following radio loss of lock between the ground trans- mitter and airborne receiver. Each transponder exciter auxiliary oscillator frequency shi_ a different amount determined by its individual characteristics and align- ment conditions. It was concluded that the effect noted on the flight radio frequency is inherent in the transponder design. A special radio test was conducted on 9 April 1965 to verify an interferometer effect noted at the spacecraft receiver as a result of dual signal paths from the high- gain and low-gain antennas when the RF circulators were switched to receive on the low-gain antenna. Results of the test indicate that the circulator loss through the nonoperating path is approximately equal to the gain of the high-gain antenna. Flight command tests were conducted on 9 and 14 June 1965 to verify the ability of the command detector in the flight command subsystem to lock up and remain in lock with various amounts of radio receiver static phase error (SPE) and with the command detector voltage-controlled oscillator (VCO) offset various amounts. In addition, the command detector was monitored for command bit errors resulting from adjustment of radio static phase error while command was in lock. No undesirable effects were detected as a result of the test conditions while displace- ments and tuning rates were held within reasonable magnitudes. An additional test was conducted on 12 July 1965 that demonstrated the ability of the flight command detector VCO to lockup to the ground command VCO while radio was adjusting the receiver SPE at the rate normally employed by the Deep Space Instrumentation Facility (DSIF) stations. Video storage count-two logic circuit tests were con- ducted on 11 June 1965 to determine the effect of power o..k .... _,_m tr_n_;ont-e Trnn_iant_ ware, generated bvo ul.po v o_u. .................. issuing cominands that were applicable to the encounter portion of the flight and by inducing main booster regulator failures and turning gyros on and off with the spacecraft battery disconnected. None of the test con- ditions were able to disturb the count-two circuits. A series of six power subsystem tests was conducted between 21 June and 12 July 1965 to determine the effects of all possible encounter load switching sequences when the spacecraft was powered by solar power at Mars levels and without a spacecraft battery to absorb current transients. The results of this series of tests indicated that encounter equipment turn-on transients could be supported by the solar panels alone and that the current surges could be reduced to still lower levels by following a procedure that reduced power requirements during encounter initiation commands. The final flight support test was conducted on 14 July 1965; it consisted of an encounter conducted on the PTM in step with the flight encounter performed by Mariner IV. The objective of this test was to provide a test specimen preconditioned to the flight mode through- out the encounter period to facilitate investigation of any nonstandard conditions that might be observed in the flight data. No questionable conditions were noted and the PTM was tan'ned off following exit from occultation by the flight spacecraft. c. SFOF encounter preparation tests. Thirteen en- counter tests were conducted for the SFOF between 12 April and 9 July 1965 to enable evaluation of various encounter command sequences and to provide operations personnel with experience in the spacecraft reactions to encounter conditions as preparation for the actual event. Since several of the tests were intended to fur- nish practice in interpreting the telemetered data from the spacecraft, special instrumentation was attached to the PTM during those tests to enable the system test per- sonnel to produce variations in temperature and signal information to telemetry and to simulate various failure modes in the encounter instruments as well as the space- craft systems that were essential to the encounter 65 JPL TECHNICAL REPORT NO. 32-881 sequence. Spacecraft operating time during this series of tests was 122.6 hr. Tests on 13, 15, and 20 April 1965 were devoted to evaluation of various command sequences that were proposed to prevent loss of Canopus acquisition during the encounter period. The commands were intended to place the spacecraft on inertial control and correct for drift with incremental commands. The procedure was successful but was subsequently ruled out as unnecessary. The objective of these tests was to enable SFOF personnel to determine the optimum procedure for opening the science cover and prepositioning the scan platform, and to practice the selected procedure and timing sequence. Several modes of operations were employed during the tests to enhance the responses to be investigated. Since transmission time is an important factor in the timing of spacecraft commands, both uplink and downlink radio transmission times were simulated during several of the tests. For each test, the spacecraft was conditioned to a pre- encounter cruise mode by the system test personnel, who then reacted to SFOF requests for commands through a procedure simulating the DSIF command links. Space- craft telemetry signals were processed to a teletype format and transmitted to SFOF. The tests demonstrated the feasibility of an early encounter to open the science cover and position the platform as well as providing the necessary practice for the SFOF personnel to develop their technique. SFOF encounter practice and backup mode tests were conducted on 27 May and 3, 4, 15, 24, and 28 June 1965. Simulated failures and nonstandard operating conditions were introduced into the telemetry signals during each of these tests to provide the operations personnel with practice in the analysis and corrective action to be used for each anomalous condition. SFOF scan position tests were conducted on 1 June and 1 July 1965. Although these two tests were con- ducted specifically to investigate the timing of the scan platform positioning commands, the technique was prac- ticed during several of the other SFOF encounter practice tests. No operational problems were introduced during these tests. SFOF operational readiness tests were conducted on 6 and 9 July 1965. Spacecraft operational problems were simulated during the first test, but the second was com- pleted with nominal operating conditions. d. Printed circuit ring harness evaluation tests. Printed circuit ring harness evaluation tests were conducted on the spacecraft from 15 March until 6 April. These tests consisted of detailed investigation of all signal and power circuits in the 9W1 and 9W9 ring harnesses while the spacecraft subsystems were connected to the original wired harnesses and while they were connected to the printed circuit ring harnesses. Two versions of the printed circuit ring harnesses were employed and, in addition, a combination of the original 9W9 wired harness for power circuits and the printed circuit harness for the signals normally routed through the 9W1 was examined. The method of investigation was to photograph all signals on the original wired harness by means of an oscilloscope camera, then repeat each of the photographs under iden- tical conditions except with the printed circuit harness substituted for the wired harness. Subsequent tests on the second printed circuit harness consisted of comparison of the oscilloscope patterns with the original photographs and repeating photographs of the ones that indicated a change. An analysis of the results of these tests appears in Ref. 5. e. Magnetic mapping and degaussing investigation. Magnetic mapping and degaussing tests were conducted from 18 to 24 May 1965 with investigations extending from 12 to 26 May 1965. The objective of these tests was to determine the stability of the spacecraft magnetic fields. The spacecraft was not operated during the mag- netic tests. f. Miscellaneous tests. Miscellaneous tests were con- ducted on the radio, science, and pyrotechnic subsystems and on the read, write, verify (RWV) and launch mode control set OSE equipments from 26 February to 18 June 1965. These tests were performed to investigate sub- system problems and obtain information about spacecraft and support equipment that, while not directly related to the Mariner IV flight, was necessary for evaluation of the program. All miscellaneous tests were conducted on a noninterference basis with the flight support tests. Total spacecraft operating time for this series of tests was 50.3 hr. Ten tests were conducted on the radio subsystem to recheck calibrations and to investigate a temperature sensitive anomaly that caused a momentary loss of 66 .uplink lock approximately 1 hr after power turn on. This problem was peculiar to the PTM radio subsystem and was the result of transponder instability confined to a narrow temperature range. The subsystem was removed from the spacecraft and returned to the laboratory for more complete testing. Science subsystem tests were conducted on the Trapped Radiation Detector, to investigate the effects of high level radiation sources; on the teievision camera, to verify that apparent voltage variations are introduced by the telemetry sampling system and are not reflected back to the television circuits; and on the data automation system, to investigate noise introduced into the Cosmic Dust Detector readout. A pyrotechnic test was conducted on 16 April 1965 using RF sources to illuminate the spacecraft and to simulate launch environment with live squibs installed in all pyrotechnic positions. No problems occurred during the test. One test was conducted to determine compatibility of the spacecraft with the launch mode control set as- sembled to operate the PTM spacecraft during vibration tests. Operation from the test set was normal. The electromagnetie interference group conducted tests to identify the RF signature of the PTM spacecraft and to investigate noise generated by the on board radio equipment and the attitude control gas valve operation. Several tests were conducted to verify ground com- mand equipment compatibility with the operating space- craft prior to shipment to the various DSIF stations. Since many of the listed tests involved individual subsystems, several were performed simultaneously when they were noninterfering. 4. From 16 July to 23 September 1965 The post-encounter phase of PTM spacecraft opera- tions was devoted primarily to completion of tests having value as sources of information for future spacecraft design. Two tests were performed to assist with interpre- tation of Mariner IV flight data, and two were performed to investigate anomalies in the PTM radio subsystem. Tests were run to obtain information on spacecraft noise-reduction techniques; there were examinations of power return and spacecraft grounds, as well as of signal return lines for noise, while various grounding techniques JPL TECHNICAL REPORT NO. 32-881 were tried. Additional photographs were taken of signals from modified printed circuit harnesses; and a final magnetic mapping was completed. Flight support tests were conducted to develop a technique to control ground transmitter carrier level for spacecraft automatic gain control (AGC) calibration, analyze the effects of encounter commands on the radio transmitter frequency, and attempt to determine the cause of ,_t-1.... _,,_a ,,."__t_._... ....;gnPJ_ observed in Mariner IV data during encounter. Radio subsystem tests were conducted to investigate an uplink lock anomaly and cavity amplifier power instability; both problems were peculiar to PTM equip- ment. One system verification test and a final systems test were performed during the period. Tests performed during this period are listed in Table 15. Table 15. Test sequence, 15 July to 23 September 1965 No. 1 7/20/65 2 7/29/65 3 7/30/65 4 8/11/65 5 8113165 6 8/16/65 7 8/19/65 8 8/24/65 9 8/25/65 10 8/26/65 11 9/8/65 12 9/9/65 13 9/13/65 14 9/21/65 Date Test Power investigation of noise on spacecraft ground, continuing through 7/28/65 OSE power ground noise investigation Magnetic mapping and degaussing, continuing through 8/9/65 System verification test, completed 8/12/65 Radio uplink lock anomaly investigation Spacecraft return system noise investigation, continuing through 8/18/65 Radio investigation of periodic amplitude varia- tions of cavity amplifier output power, completed 8/20/65 Printed circuit ring harness investigation, completed 8/25/65 Radio investigation of ground transmitter power suppression techniques Coupled noise investigation, continuing through 8/31/65 Printed circuit ring harness investigation Video storage-data automation system interface noise investigation, completed 9/10/65 Final system test, continuing through 9/15/65 Radio encounter investigation, completed 9/23/65 6"7 JPL TECHNICAL REPORT NO. 32-881 a. Power subsystem noise investigation. Power sub- system noise investigations were carried out during 20-29 July, 16-18 August, and 26-31 August 1965. Total spacecraft operating time for these tests was 25.1 hr. Both spacecraft and OSE power and signal returns were examined for excess noise and circulating currents. The possibility of reducing noise induced into the signal lines to data encoder by making a hard ground connec- tion between the spacecraft power return and spacecraft was tested by connecting a fuse between the two points in question. The first attempts to connect the fuse re- suited in abnormal spacecraft operation and high noise levels in all signal circuits, but the problem was resolved as a wiring error resulting from an engineering change instruction that was not worked on the PTM power distribution subassembly module. The problem had not been noticed previously due to the isolation between primary power and the spacecraft bus. After completion of the applicable engineering change instruction (ECI), no problems were noted when the spacecraft power return and spacecraft frame connected. An attempt was also made to compare the susceptibility of the spacecraft subsystems to power transients with the power return floating as in the normal configuration and with it con- nected to the spacecraft frame. No change in response to power transients was noted. OSE power distribution return lines were examined for circulating currents that could cause induced noises and the result was the location of a faulty transformer in the science data format converter. The faulty transformer was disconnected from the system and both return currents and noise were reduced within normal levels. Tests were also conducted to determine the effects of various shield connections and ground schemes, but no definite improvements were achieved. It was determined that noise from signals on one line can be induced into another line by a common conducting cover or shorted loop encircling them. b. Magnetic mapping. Magnetic mapping of the space- craft was conducted during the period 30 July through 9 August 1965 to determine what effects were produced in the spacecraft magnetic field by the electrical tests conducted since the mapping tests completed 26 May 1965. The following changes were observed in the perma- nent magnetic field (P): Px from -33/4 to -41/4 _, Py from 0 to 1/2 _, Pz from 4V4 to 53A p The spacecraft was not operated during the magnetic tests. c. System verification test. A system verification test was conducted on 11 and 12 August to verify that space- craft operation was normal following the mechanical work necessitated by the magnetic mapping procedure. Spacecraft operating time for this test was 7.7 hr. All spacecraft operation was normal. d. Radio special tests. Radio investigation of an anom- aly in the uplink lock circuitry was conducted on 13 August 1965 to determine if removal and reinstallation of the radio subsystem had caused a change in previously noted conditions. The anomaly was still present and unchanged. Spacecraft operating time for this test was 2.8 hr. Radio investigation of power variation from the cavity amplifier was conducted on 19 and 20 August 1965. The power variation was found to be of small amplitude and to have a period of approximately 35 min. It was attributed to the temperature compensating features of the cavity amplifier and was not of importance. Space- craft operating time for this investigation was 4.0 hr. e. Printed circuit ring harness investigations. Printed circuit harness tests were conducted 24 and 25 August and 8 September 1965. Objective of these tests was to determine if noise coupled between circuits could be reduced by grouping noise producing lines away from low level signal circuits and to determine the effects of removing transpositions from the power lines. Very little change in noise levels was observed. Spacecraft operating time during these tests was 4.7 hr. f. Mariner IV flight support tests. Three tests were conducted in direct support of the Mariner IV flight data analysis effort. Spacecraft operating time for these tests was 8.8 hr in addition to the radio carrier level suppres- sion tests which were conducted concurrently with the printed circuit ring harness tests. Radio tests were conducted on 24 and 25 August 1965 to develop and verify a method to control uplink carrier level suppression by means of a 10-kc modulating signal applied to the ground transmitter. The objective of this method is to enable precise control of the uplink signal level transmitted from the DSIF stations in order to calibrate the Mariner IV receiver AGC voltage during flight. The technique was successful and achieved carrier level suppression of 20 db. Since the modulating 68 JPL TECHNICAL REPORT NO. 32-881 w frequency was 10 kc the sidebands that it generated were outside the spacecraft receiver passband. Video storage/DAS interface noise tests were con- ducted on 9 and 10 September 1965. The test objective was to determine ff false video storage end-of-tape signals can be issued to DAS by noise spikes produced by tape start-stop commands. Start-stop commands were issued to video storage first from the OSE through the direct access lines and then from DAS by conducting an encounter. Both methods produced noise spikes of ap- proximately equal amplitude on the video storage/DAS end-of-tape line, but DAS only issued false end-of-tape signals to data encoder when the start-stop commands were issued from the OSE. Since a much greater number of start-stop commands were issued by the OSE method than by encounter conditions, it is possible that a greater number of encounter start-stop commands may have resulted in a false end-of-tape signal from DAS to data encoder. Radio encounter tests were conducted on 21 and 23 September 1965 to determine ff the encounter sequence produces a change in the transponder transfer command voltage. The object of this test was to look for an explana- tion for a slight frequency shift noted in the Mariner IV downlink signal following the planet encounter and the dark encounter. The results of the test indicated that the transfer command voltage does change ff the encoun- ter is controlled by radio commands, but does not change ff the encounter sequence is initiated and terminated by CC&S commands without radio using two-way lock. The tests were not able to determine ff the effect is the result of transferring from auxiliary oscillator to the VCO and back to auxiliary oscillator or the result of the command source. g. Final system test. A system test was conducted from 13 through 15 September 1965 to assure spacecraft integrity prior to storage. Spacecraft operating time was 13.3 hr. All spacecraft subsystems functioned normally and the test was considered satisfactory. 5. Spacecraft Configuration for Storage The PTM spacecraft has been prepared for storage and testing has been terminated. Table 16 describes the detailed configuration of the PTM spacecraft at the time of its commitment to storage. Also included is such spacecraft-related information as planetary instrument alignment angles, spacecraft behavior idiosyncrasies, transmitter frequencies, and nominal spacecraft power demands in various power modes. Table 16. PTM configuration and nominal characteristics at time of storage a. Nominal spacecraft power demands Po_fer Mode Spacecraft configuration demand, _ w 171 bPrelaunch science checks Prelaunch Launch i Separatlon Sun acquired (Earth cruise with gyros on) Earth cruise (Canopus acquired) Midcourse lO0-day cruise Mars cruise Mars cruise with gyros on Encounter Playback Cruise science on, encounter science off, RF power down (cavity), OSE relay hold on, gyros on, external power All science off, RF power down (cavity), OSE relay hold on, gyros on, external power All science off, RF power down (cavity), spacecraft relay hold on, gyros, on, internal power. VS$ not in launch mode Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros on, internal power Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros on, solar power Earth mode Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros off, solar power Earth mode Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros on, internal power Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros off, solar power 100-day Cruise science on, encounter science off, RF power up (cavity), t;elay hold off, gyros off, solar power Mars mode Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros on, solar power Mars mode Cruise science on, encounter science on, RF power up (cavity), relay hold off, gyros off, solar power Mars mode Cruise science off, encounter science off, RF power up (cavity), relay hold off, gyros off, solar power Mars mode, video storage 2.4-kc power on 149.5 b 155 b 189 b 195 b 140 b 198 b 140 b 142 b 196 b 173 b 123 c aThe demand of the battery charger is variable and must be added to the figures shown above to obtain total spacecraft demand. The demand of the battery charger is 27w when charging at the maximum rate of 300 ma with 44-v input to the charger. The power demands shown are based on the averages of system tests 3, 4, and AFETR I and 2 and the space simulator test. bThe TWT uses 27w more than the value shown with the cavity on. cMeasurement was made with TWT on, then 27w subtracted to provide nomi- nal figure for cavity. 69 JPL TECHNICAL REPORTNO. 32-88! Table16a.(Cant'd) Mode Spacecraft configuration Power demand, a w 144 _Playback with cruise science Failed booster with gyros on Failed booster with gyros off Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros off, solar power Mars mode, video storage 2.4-kc power on. Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros on, solar power 100-day mode. Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros off, solar power 100-day mode. 189 b 142 b aThe demand of the battery charger is variable and must be added to the figures shown above to obtain total spacecraft demand. The demand of the battery charger is 27w when charging at the maximum rate of 300 ma with 44-v input to the charger. The power demands shown are based an the overages of system tests 3, 4, and AFETR 1 and 2 and the space simulator test. bThe TWT uses 27w more than the value shown with the cavity on. _Measurement was made with TWT on, then 27w subtracted to provide nomi- nal figure for cavity. b. Nominal planetary instrument alignment Instrument Measurement, deg Fields of view Television Scan sensor Narrow-Angle Mars Gate Look angles Television (reference) Scan sensor Narrow-Angle Mars Gate 1.1 X 1.1 48.5, cone 1.5 X 2.5 Clock angle Cone angle O.00 1.00 0.10 120.00 119.70 120.10 c. Nominal planetary scan platform alignment Parameter Measurement Clock angle (pinned) Backlash Voltage indication in stowed (pinned) position 238 deg 02 min arc 21.47 min arc 162 ±1.9022 v d. Nominal Canopus sensor alignment Parameter Measurement Clock angle 259 deg cw from Bay II centerline Table 16. (Cant'd) e. Radio frequencies Classification Frequency, Mc Nominal 25°C uplink frequency for VCO at zero SPE Nominal 25°C downlink frequency for VCO at zero SPE Nominal 25°C frequency for one-way transmission Exciter A Exciter B 2115.698747 2297.591400 2297.586552 2297.592936 f. Nominal radio Classification RF power up Power amplifier A Power amplifier B RF power down Power amplifier B )ower output Power, dbm 40.0 38.0 28.0 g. Permanent magnetic output a Magnetometer axis Measurement, ")' X Y Z -- 4.25 +0.5 + 5.75 aSpacecraft perm without solar panels. h. Video storage record time Period Time, sec During encounter 1512 _ 1.0 i. Spacecraft idiosyncrasies Subsystem Idiosyncrasy Radio Command Data encoder Possible loss of lock on command DC-9 at weak signal level Loss of lock on command DC-9 Power transients occasionally causing medium deck skips F. MC-4 Spacecraft Operations and Disposition 1. From 30 September 1964 to 9 April 1965 (at AFETR) Mariner spacecraft MC-4 arrived at AFETR on 29 September 1964. Unloading and assembly operations 7O werecompletedon 2 October 1964. System level tests were performed as necessary to assure that the space- craft was in flight-worthy condition to support the launches of MC-2 and MC-3 as spares or as a replace- meat. Following the launch of MC-2, two updated TA solar panels were added to bring the spacecraft to launch configuration as standby for MC-3. On 1 December 1964 the spacecraft was placed in temporary storage at AFETR until 9 April 1965. The tests conducted at AFETR are listed in Table 17. Table 17. Operations and tests of MC-4 at AFETR Operation or test Date Assembly System verification test System test AFETR No. 1 Autopilot test Final inspection, disassembly and assembly 'System verification test Attitude control leak check System verification test Science subsystem alignment (SPINAl.) check Science calibration checks System test AFETR No. 2 Current loop magnetic mapping test Magnetic mapping Attitude control nozzle flow rate test Final mechanical preparation 1o/2/64 1o/2/64 io/2/64 to 10/6/64 10/6/64 lo/B/64 to lo/15/64 lO/16/64 io/2g/64 to1i/2/64 1i/1o/64 11/io/64 11112164 11/13/64 11/17/64 11/17/64 to 11/18/64 11/18/64 11/18/64 11/19/64 I,o 11/24/64 a. Spacecraft arrival. The spacecraft arrived at AFETR on 29 September 1964. Unloading and mechanical assembly were started on 30 September and completed 2 October 1964. All units were flight qualified except the science Cosmic Dust and Trapped Radiation Detectors, which were PTM modules. The flight units for those experiments did not arrive at AFETR prior to spacecraft assembly. A science DAS unit scheduled for use in MC-3 was installed in MC-4 to undergo the system test. JPL TECHNICAL. REPORT NO. 32-881 b. System verification test. Spacecraft testing started on 20 October 1964 with the system verification test. Spacecraft operating time was 8.2 hr. No significant problerfis-were encountered but the pitch gyro output was noisy; it was evaluated during succeeding tests. c. System test AFETR No. I. System test AFETR No. 1 was started 2 October and completed 6 October 1964. Spacecraft operating time for the entire test was _qn.6 hr. To _s_t v_th sehed,_,|ing_ the test was conducted in segments during a 5-day period. Several significant problems occurred and were corrected: . The Canopus sensor shutter solenoid failed to operate because the wires were not properly con- nected or soldered. During an investigation into the solenoid problem, the wires were shorted, causing the main booster regulator to drop out. Subsystem laboratory tests determined that no damage had resulted from the short cireut. All wiring was re- connected and soldered properly. . The pitch gyro channel again produced excessive noise and the entire gyro package was changed to correct the problem. . No track change was obtained on the video storage tape recorder at end of tape or by OSE commands. The track change relay had been hung up by an OSE operator error prior to test start. The relay was released in the laboratory and normal operation restored. d. Attitude control autopilot test. An attitude control subsystem autopflot test was conducted on 6 October 1964. Spacecraft operating time was 1.8 hr. No significant problems were encountered. e. Final inspection. Spacecraft disassembly for final inspection was started on 8 October 1964. Final inspection and reassembly were completed on 15 October 1964. Final inspection of each spacecraft is normally performed at JPL, Pasadena, prior to shipment to AFETR, but in the case of MC-4, the inspection was delayed until the spacecraft had been delivered to AFETR to demonstrate feasibility of performing the inspection at that facility. f. System verification test. A system verification test was conducted on 16 October 1964. Spacecraft operating time was 10.4 hr. The purpose of this test was to verify 71 JPL TECHNICAL REPORT NO. 32-881 that the spacecraft system functioned within specifica- tions following the final inspection and reassembly. Two problems were recorded: 1. Scan platform unlatch failed to give a data encoder event. 2. The science Ionization Chamber and Cosmic Ray Telescope did not calibrate properly and were re- moved for test and calibration. g. Attitude control gas leak check. An attitude control gas leak test was started on 28 October and completed 2 November 1964. The system was within specifications. Following the leak test, the spacecraft was assembled in complete flight configuration and placed on standby to support the launch of MC-2. h. System test AFETR No. 2. System Test AFETR No. 2 was started and completed through half of the first encounter on 10 November 1964 as a verification of spacecraft performance and as the first section of the final spacecraft system test. Spacecraft operating time was 6.7 hr. Significant problems were Cosmic Ray Tele- scope noise and failure of the scan platform pin release to give a data encoder register event. i. Science subsystem alignment check. Science subsys- tem alignment (SPINAL) checks were conducted on 10 November 1964. Spacecraft operating time was 2.5 hr. j. Science calibration checks. Science subsystem cali- brat'ion was started on 11 November and completed on 12 November 1964. Spacecraft operating time was 16.6 hr. No significant problems were encountered; how- ever, one previously observed problem recurred when the removal of the science scan platform pin failed to produce a data encoder event. k. System test AFETR No. 2 (Concluded). All remain- ing portions of the final system test started on 10 Novem- ber were conducted during the period 13 to 17 November 1964 to verify that the spacecraft was ready for launch if needed as a replacement for MC-3. Spacecraft operat- ing time for the test was 34.8 hr. Two significant problems were observed and corrected during the test. 1. The Cosmic Ray Telescope noise problem was at- tenuated by the addition of an extra cover screw to tighten the cover and complete the shielding con- tinuity. 2. A circuit board in the scan platform unlatch event circuit was replaced to correct the intermittent e_zent readout. l. Magnetic mapping current loop test. A magnetic mapping current loop test ran from 14--15 November. Spacecraft operating time was 4.4 hr. Results of the test were normal as compared with the information obtained during tests of MC-2 and MC-3. m. Magnetic mapplng. Magnetic mapping of the spacecraft was conducted on 18 November 1964. The spacecraft was not turned on during this test. n. Attitude control nozzle flow rate test. An attitude control nozzle flow rate test was conducted on 18 No- vember 1964 under the direction of the attitude control cognizant engineer. o. Final mechanical preparation. Final mechanical preparation of the spacecraft for launch was started 19 November and completed 24 November 1964. Four flight solar panels were installed including two that had been updated from type-approval to flight configuration. The spacecraft was placed on standby as a launch replacement for MC-3 on 29 and 30 November 1964. On 1 December 1964 the spacecraft was placed in storage under a continuous nitrogen purge in Building AO at AFETR, where it remained until 9 April 1965. The space- craft test history from 25 November 1964 to 24 May 1965 is recorded in Fig. 52. 2. From 9 April to 15 June 1965 (at JPL) On 9 April 1965 the spacecraft was removed from stor- age at AFETR, shipped to JPL, Pasadena, subjected to verification and system tests, operated in support of Space Flight Operations Facility encounter practice tests, then finally, placed in temporary storage in the SAF on 15 June 1965. All test results were within specified values. The tests conducted at JPL prior to storage of the spacecraft are listed in Table 18. a. Return to ]PL. The spacecraft was removed from storage at AFETR and shipped to JPL, Pasadena, on 9 April 1965. Following arrival on 14 April, mechanical inspection and preparation of the spacecraft for system test was completed. It was connected to the system test complex on 28 April 1965. 72 MARINER C-4 TEST AND OPERATIONS SUMMARy DATE 11/25/64 to 419165 419/65 4113165 4130165 413O165 III 514165 I II 15/5/65 I I, 5/6/65 111 II 5111164 5113165 5I]4/65 5118165 5119165 5/zo/65 5/21/65 l 5/24/65 SPACECRAFT OPE RATIONS SYSTEM TEST NO. HOURS OF SIC OPERATIONS OPERATIONS 480 490 500 510 520 530 540 550 560 TOTAL SIC OPERATING TIME 540.6 HOURS CRUISE SCIENCE AND CALIBRATE PORTIONS OF SYSTEM CM 300.03 (10.0 HRS) ATTITUDE CONTROL BACKUP COMMAND VERIFICATION AND SOLAR PRESSURE VANE ACTUATOR VERIFICATION, RADIO BA( COMMAND VERIFICATION AND SUBSYSTEM VERIFICATION POR OF SYSTEM TEST CM B00.03 (3.4 HOURS| GYRO, AUTOPILOT, AND SUN SENSOR VERIFICATION PORTIONS SYSTEM TEST CM 300.03 (0.8 HRS| CC&S COUNTER TESTS (EXCEPT MIDCOURSE] FOR SYSTEM TEST CM 300.03 (6.3 HRS} PYRO SPECIAL TEST AND CC&S MIDCOURSE PORTION OF SYSTEM TEST PROCEDURE CM 300.03 (1.6 HRS) EMI RF SIGNATURE (I.0 HRS) f SYSTEM VERIFICATION TEST CM 301.01 (7.1 HRS) ARRIVED JPL REMOVED FROM STORAGE & SHIPPED TO JPL S/C IN STORAGE AT ETR JPL TECHNICAL REPORT NO. 32-881 TEST ,KUP Jrl J SF'OF ENCOUNTER PRACTICE TEST (5.3 HRS) ' SIeOF ENCOUNTER PRACTICE TEST (7.0 HRS) POWER SPECIAL BOOSTER REGULATOR TEST (l. 8 HRS) Fig. 52. MC-4 test history, 25 November 1964 to 24 May 1965 JPL TECHNICAL REPORT NO. 32-881 Table 18. Tests at JPL prior to storage Title Date System verification test Radio frequency signature System test Power booster regulator investigation SFOF encounter practice test SFOF encounter sequence logic flow diagram checkout Special cleanup tests 4/29/65 to4/3o/_ 4/3o/_ 51416s to 5/2o/_ 51v316s 5/141_s s/lg/_s 5121165 b. System Verification Test and RF Signature Investi- gation. A system verification test was started on 29 April and completed 30 April 1965. Spacecraft operating time was 8.1 hr. The purpose of this test was to verify that the spacecraft had not sustained damage or deterioration during storage and shipment and that it was still com- patible with the PTM system test complex. All spacecraft systems functioned normally and all apparent discrep- ancies were resolved as being external to the spacecraft. Concurrently with the system verification test, an RF signature investigation was conducted by the EMI group. This investigation provided a knowledge of spacecraft radiation characteristics in the HF, VHF, and UHF por- tions of the RF spectrum, and was a continuation of a similar test conducted on the PTM. Since the spacecraft was to be operated for several weeks to support the flight of Mariner IV, the television optics were replaced with the units from the PTM to avoid wear on the shutter in the flight qualified unit. c. System Test. A complete system test was performed from 4 to 20 May 1965. Since continuous operation was not practical, the test procedure was divided into con- venient segments that were conducted on separate days when they best fitted the schedule of tests required to support the Mariner IV flight. The spacecraft was oper- ated a total of 34.5 hr for the entire system test. Objective of the system test was to provide a complete system operational checkout prior to spacecraft storage, to guarantee operational readiness of a requirement for spacecraft operation if generated in the future. The system test resulted in the conclusion that all subsystems and the entire spacecraft system were operational. During the test, several apparent anomalies were ob- served, but all were resolved as the result of non-standard conditions or support equipment problems. One incident of particular interest was the result of performing a main booster regulator failure exercise while the spacecraft was operating from external power, rather than simulated solar power. The transfer resulted in a cruise science turnoff. Normally the failure procedure is performed under flight conditions with the spacecraft battery con- nected in the circuit to supply the current demand and no exeept_ion_l transients are noted. The test was re- peated with the spacecraft operating from simulated solar power; all results were according to procedure. d. Power subsystem special booster regulator test. A special power subsystem main booster regulator failure test was conducted on 13 May 1965, with the spacecraft operating from the various power sources available in the system test configuration. Spacecraft operating time was 1.8 hr. The purpose of the test was to verify the effects observed during system test when the main booster regulator test was performed with the spacecraft operaHaag from external power. Simulated main booster regulator failures were in- duced while the spacecraft was operating on external power with results identical to those observed during the system test. The test was repeated with the spacecraft operating from simulated solar power and the results were the same as those normally observed during the standard system test procedure. Since the external power source is current-limited and no spacecraft battery is connected in the circuit during the external power mode of operation, no firm power source is available to supply the switching transient de- veloped by the changeover from main to maneuver booster regulator. Operation from the simulated solar power sources is normally performed with the spacecraft battery connected in flight configuration, to provide a high current source to supply the switching transient, so very little disturbance is noticed in the spacecraft sub- systems at the time of changeover from main to maneuver booster regulator. Spacecraft operation was considered normal for the test configuration involved. e. SFOF encounter practice tests. Two SFOF encoun- ter practice tests were conducted to support the flight spacecraft (Mariner IV) during the interval that the PTM was committed to magnetic mapping and degaussing. 75 JPL TECHNICAL REPORT NO. 32-881 The first test was conducted on 14 May 1965 resulting in 7 hr of spacecraft operating time. Since the test objec- tive was to furnish SFOF with spacecraft telemetry sig- nals similar to those expected from Mariner IV during encounter, the test spacecraft was conditioned as closely as possible to the flight mode of operation without mak- ing changes to MC-4. This procedure generated some attitude control and science indications similar to the flight telemetry but did not provide flight temperature readings. Several minor problems were introduced into the telemetry data to simulate possible flight problems to make the exercise more meaningful. All of the indications" were produced without resorting to special equipment or changes in the MC-4 spacecraft. The test resulted in one P/FR 5 written against the radio subsystem because the telemetry indication of transmitter power output varied continually throughout the test. No corresponding variations were detectable in direct access measurements however, and later tests on 21 May 1965 resolved the problem as an incorrect ground on the recorder in attitude control OSE. 5P/FRs initiated at SAF are listed in Table 19. 28.131 cg TO JET VANE CENTER (POINT OF ROTATION OF THRUST VECTOR) AT 2 deg TO HORIZONTAL I- 0.500 MAXIMUM OFFSET CAPABILITY ;0.70711(X-_)=0.5065 _ ;INITIAL KNOWN OFFSET= 0.231 --_ PiPS ¢. _'--_ -/ JET VANE CENTER = -12.274 +Z I ------:C_¢ / _ (z L MOTOR TILT ANGLE TR_NS,AT,ON,N "_-__ _-----Z :-,._ MOUNTINGPLANE:a,O_-_I-'_ \ MOUNT,N0_LANEO_P,PS--_ _-P,VOTPO,NT 21.916 Fig. 53. MC-4 PIPS positioning 76 JPL TECHNICAL REPORT NO. 32-881 Table19.SAFinitiatedP/FRsagainstMC-4spacecraft Problem Date open Date resolved Date dosed ResolutionSubsystem Structure Radio Power CC&S Ref 4101303 4101303 E-A1 E-A7 02 02 02 02 Data 6A Enceder O2 O2 4A 13 4A11 4A3 04/OSE O4 5A SAFP/FR! No. 2 Leg C thread bid in insert 4 Two screws frozen in insert, Legs B end F 5 Loose insert •vo ..'_=C.. m'_-_,._..uo ! ,4 ,emn ° 18 Malfunction in system 22 Malfunction in releasing boost inhibit 25 Malfunction, lost lock 26 Malfunction, in mixer preamplifier 57 Noise in receiver 80 Data not varied -- 105 -- ] 09 16 Data encoder lost lock 19 Bed insert, screw binds 33 TA panel seemed to chatter, Leg D, Bay III 37 Noise malfunction 81 Malfunction in system 48 Power supply in central timer burned up power transformer (CC&S) 20 Data encoder trouble stay- ing in sync while homing 7/21/64 4129164 8p/64 8/12/64 8/13/64 8/13/64 9/3/64 5/14/64 8/5/64 8/7/64 8/23/64 8/26/64 5/13/65 9/11/64 8/7/64 The second SFOF test was conducted on 18 May 1965, resulting in 5.3 hr of spacecraft operation. The objective of this test was to support SFOF checkout of their pro- posed encounter sequence logic flow diagram. As in the previous test, the spacecraft was conditioned to the pre- encounter cruise configuration; commands were then executed as directed by SFOF. All spacecraft operations were normal. f. Spedal cleanup tests. Three special tests were con- ducted on 21 May 1965. Spacecraft operating time was 2.7 hr. The objective of the tests was to ensure that all problems were correctly explained prior to spacecraft storage. First, a special test was conducted on the video storage subsystem to verify an explanalJon by the operator of 8/3/64 4/29/64 8/20/64 8/2O/64 10/I/64 6/8163 8/25/64 8/12/64 9/I I/64 9/29/64 9/21/64 8/14/64 7/28/64 8/21/64 8/6/64 7/21/63 8/20/64 S/20/64 9/19/64 10/I/64 10/8/64 9/7/65 8/31/64 8/18/64 9122164 1o/i/64 10/1/64 $125164 Press-nut replaced Reworked, correct HiTi screws used Insert removed, new insert installed Replaced Replaced cable corrected problem Reworked per ECR 7636, 7702 Not applicable Replaced Not applicable Panel updated Telemetry calibration procedure changed Insert replaced Use restricted to non-flight None (OSE or cabling) Still open Power supply replaced None the circumstances that resulted in loss of part of the recorded video during system test. The test established that the video was inadvertently erased by the video storage operator when he advanced the tape prior to the encounter science turn-off command. Since this is an OSE function, it cannot happen to the flight spacecraft. The second test was an investigation of the radio sub- system to determine the apparent variation of transmit- ted power as indicated by telemetry during the SFOF test. An abnormal ground in the attitude control OSE rack was the result of a jumper that had been used to connect the attitude control strip chart recorder to the OSE cabinet. The jumper closed a ground loop involving spacecraft ground with the OSE ground system and introducing noise in the spacecraft data encoder input. The noise amplitude was sufficient to disturb the 100 mv telemetry channels from radio. Examination of the 77 JPL TECHNICAL REPORT NO. 32-881 previous test data revealed corresponding variations in all of the radio 100 mv channels. Again the cause of the problem was the result of support equipment and is not a flight problem. The third test was a repeat of the power main booster regulator failure test performed on 13 May 1965. Its purpose was to provide photographic documentation of the transients observed during the earlier test. All results were the same as obtained previously. As a result of the test sequence, the spacecraft was considered to be in flight condition and to have suffered no deleterious effects from storage at AFETR or ship- ment to JPL. g. Spacecraft Preparation For Storage. Preparations for storage were completed during the period 25 May to 15 June 1965. All components requiring special service prior to storage were properly prepared and all items requiring separate storage containers were removed. The spacecraft was placed in its storage container under Table 20. MC-4 configuration and nominal characteristics at time of storage a. Nominal spacecraft power demands Mode Spacecraft configuration Power demand, w" 168 bPrelaunch science checks Prelaunch Launch Separation Sun acquired (Earth cruise with gyros on) Earth cruise (Canopus acquired) Earth cruise (minimum demand) Midcourse Cruise science on, encounter science off, RF power down (cavity), OSE relay hold on, gyros on, external power All science off, RF power down (cavity), OSE relay hold on, gyros on, external powel All science off, RF power down (cavity), spacecraft relay hold on, gyros on, internal power Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros on internal power Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros on, solar power Earth mode Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros off, solar power Earth mode All science off, RF power up (cavity), relay hold off, gyros off, solar power Earth modt Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros on, internal power 149 b 153 b Power Mode Spacecraft configuration demand, w' lO0-day cruise 142 b i Mars cruise Mars cruise with gyros on 185 b Encounter 195 b Playback 142 b Playback with cruise science on 117 b Failed booster with gyros on 196 b Failed booster with gyros off Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros off, solar power lO0-day mode Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros off, solar power Mars mode Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros on, solar power Mars mode Cruise science on, encounter science on, RF power up (cavity), relay hold off, gyros off, solar power Mars mode Cruise science off, encounter science off, RF power up (cavity), relay hold off, gyros off, solar power Mars mode, video storage 2.4-kc power on Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros off, solar power Mars mode, video storage 2.4-kc power on Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros on, solar power lO0-day mode Cruise science on, encounter science off, RF power up (cavity), relay hold off, gyros off, solar power lO0-day mode 142 b 196 b 174 b 129 _ 147 e 186 b 144 b aThe demand of the battery charger is variable and must be added to the figures shown above to obtain total spacecraft demand. The demand of the batter charger ia 27w when charging at the maximum rate of 300 ma with 44-v input to the charger. The power demands shown are based on the averages of system tests 2, 3, 4 and 5, and the space simulator test. bThe TWT uses 27w more than the value shown with the cavity on. eMeasurement was made with TWT on, then 27w subtracted to provide nominal figure for cavity. 78 JPL TECHNICAL REPORT NO. 32-881 continuous nitrogen purge in a temporary storage area at the JPL Spacecraft Assembly Facility. Table 20 describes the detailed configuration of the MC-4 spacecraft at the _ne of its commitment to stor- age. Contained here is such spacecraft-related informa- tion as nominal planetary instrument alignment angles, spacecraft behavior idiosyncrasies, nominal transmitter frequencies, and nominal spacecraft power demand in various power modes. Table 20. (Cont'd) b. Nominal planetary instrument alignment I Instrument ........ J-- Fields of view Television 1.1 X 1.1 Scan senso¢ 50 cone Narrow-Angle Mars Gate 1.5 X 2.5 Look angles Television (reference) SCan sensor Narrow-Angle Mars Gate Clock angle Cone angle 0.00 120.00 + 0.95 120.30 --0.10 120.05 f. Nominal radio power output Classification Measurements, dbm RF power up Exciter A, power amplifier A Exciter B, power amplifier A Exciter A, power amplifier B Exciter B, power amplifier B RF power down Exciter A, power amplifier B Exciter B, power amplifier B 39.3 39.7 38.0 38.4 27.7 28.0 c. Nominal planetary scan platform alignment Parameter Measurement Planetary scan platform clock angle (pinned) Backlash Voltage ._nd._cet!on "_,n__towed (pinned) position Driving cow DHving cw 239 deg 32 rain arc 16.9 rain arc 1.8995 v 1.8886 v d. Nominal Canopus sensor alignment g. Permanent magnetic field Magnetometer axis Measurement, _' +21.0 + 6.0 "t- 32.0 h. Video storage record time Period Time, sac Total recording time durlng encounter 1512 ±1.0 i. Spacecraft idiosyncrasies Item Effect Parameter Clock angle Measurement 259 deg cw from Bay II centerline e. Radio frequencies r-Qctor Nominal 250C uplink frequency for VCO at zero SPE Nominal 25°C downlink frequency for VCO at zero 5PE Nominal 25°C frequency for one-way transmission Exciter A Exciter B Frequency, Mc 2116.724077 2298.704880 2298.680770 2298.719602 Telecommunications Flight command subsystem Data encoder Telemetry channel 302 (exciter No. 2 voltage) Has momentary loss of lock with DC-9 (ranging Oe)o Has tendency either to reset a deck or to skip a deck position with any of the commands that cause power transients (such as DC-7, DC-25, DC-26, and DC-28). The lower decks are most susceptible, but high deck may be affected. Change in spacecraft mode between two-way lock (VCO mode) and one-way lock (auxiliary oscillator mode) necessitates change in telemetry calibration curves. 7g JPL TECHNICAL REPORT NO. 32-881 Table20. (Cant'd) i. (Cant'd) Item Effect Telemetry channel 210 Guidance and control Telemetry channel 303 (Canopus cone angle) Receiver local-oscillator drive will increase with decreasing temperatures, and vice versa. When DC-13 or L-2 (but not L-3) is in effect and Canopus sensol; power is off, data number changes can be expected. j. Autopilot (nominal) Item Performance Jet vane position null offsets Jet vane actuator friction Gains determined by autopilot mix resistors Pitch and yaw channels ___ 5 deg Within tolerance Within tolerance k. Weight, center of gravity, and post-injection propulsion system (PIPS) positioning Factor Measurement Spacecraft weight with PIPS motor 57S.61 Ib Center of gravity" X Y Z Jet vane center of lift b (with motors as shown in Fig. 18) X jr YJv Z jr Motor tilt angle +0.717 in. --0.001 in. -- 12.274 in. 20.591 in. 19.884 in. -- 11.292 in. 200'00 " from horiz *MC-4 did not have four solar panels assigned to it; two flight spare panels, serial numbers 14 and 15 (substrate numbers 32 and 33) and two TA panels, Some extra b-axis gain in both channels; resulting pitch/yaw cross-coupling coefficient < ]5% serial numbers 3 and 4 (substrote numbers 18 and 14) were installed. The measured properties shown are for the motor position as in Fig. 19. hAs in the case of MC-2 and MC-3, the undeflected thrust vector misses the cg because the horizontal cg offset in the plane of Bay II is greater than the adjustment capability. V. END-OF-MISSION PLANNING A. Proposal for 1967 Reacquisition 1. Introduction The Project goals specified in the Project Development Plan dated 14 February 1964 have been met. This part of the mission, which was terminated on 1 October 1965, has been designated as Phase I. Phase II of the mission will consist of periodic contact with the Mariner IV spacecraft by the research and development facilities of the Deep Space Network at Goldstone, using advanced communications techniques. During this period, com- mands restricted to updating the Canopus cone angle (Table 21) may be sent to the spacecraft if, and when, the system capability is available. It is not expected that telemetry reception capability will be available. As the opportunity presents itself, two-way experiments will also be attempted, using the 100 kw transmitter at the 85-foot antenna site (Venus DSIF 13), in which some doppler radar information may be extracted. The project will be informed ff the DSN has been able to establish whether or not the Mariner spacecraft is transmitting. Furthermore, during this phase additional trajectory data may be provided that will yield further information on the inherent accuracy of the DSN, and which may pos- sibly improve some astrodynamical constants. Finally, the Mariner spacecraft will act as a far-field calibration device for the DSN. Phase II began on 1 October 1965. The termination of Phase II is defined as that time when the Mariner IV telemetry can again be received by the standard config- uration of the DSN. It is expected that this will occur during the first half of calendar year 1967. The exact date depends on the operational mode chosen (omnidi- rectional vs high-gain antenna mode) and on the DSN performance capabilities available at that date. 80 JPL TECHNICAL REPORT NO. 32-881 Table 21. Canopus sensor position and command sequence for 1966 and 1967 Date Position Command 1966 26 JAN- 9MAR 1 MAR-28 MAR 23 MAR-16 APR 11 APR- 7MAY 3 MAY-29 MAY OPT-2(MT-4) MT-3 MT-2 MT-1 PRESET 1 DC-17 5 DC-17 7 DC-17 7 DC-17 7 DC-17 25SEP - 4 NOV 25 OCT- 5 DEC 25 NOV- 5 JAN 1967 27 DEC -10 FEB 30 JAN- I APR 15 AUG-28 SEP 16 SEP -17 OCT 13 OCT- 8 NOV 3 NOV-30 NOV 17 NOV-20 DEC MT-I MT-2 MT-3 MT-4 OPT-1 0 PT-2 MT-3 MT-2 MT-1 PRESET I DC-17 1 DC-17 I DC-17 1 DC-17 1 DC-17 1 DC-17 S DC-17 7 DC-17 7 DC-17 7 DC-17 During Ph__se II it is hoped that significant measure- ments can be made with the RF emissions from the spacecraft as it passes behind the solar corona during conjunction. This event is due to occur in April 1966. Needless to say, sudden and continuous silence of the spacecraft during Phase II would obviate the need to proceed with Phase III. Phase III is defined as that period when normal telem- etry reception (and processing) is again possible during the reacquisition period. Phase III must, of course, in- clude those tasks or functions necessary to prepare for the actual operational period. Any version of Phase III holds the promise of an even more remarkable feat than Phase II. Starting about February 1, 1967, the trajectory of Mariner IV is such that telemetry should again be available through the use of the spacecraft high-gain antenna. At this time, how- ever, Canopus will not be in a favorable locale, and the orientation of the high-gain antenna toward Earth will have to be maintained through the use of inertial roll control This would require an operational plan discour- agingly more complex than those for Phases I and II. 2. Objeetlves But the basic opportunity is unique in that it presents the possibility of also going to a locale in space that has not yet been explored. The plans offer the possibility of coordinating results with the Pioneer missions. In addi- tion, some very valuable information as to the longevity in ._paee_ of our present technology can be obtained. a. Performance degradation measurement. With the relatively few measurements available in the Mariner IV system, the effects of performance degradation on flight equipment from extended exposure to space environ- ments are difficult to interpret, and the ability to pin- point a failed element is tenuous, at best. Nonetheless, some new information regarding longevity characteristics can be extracted from a continuing flight, and this infor- marion should be valuable to spacecraft engineering technology. A spacecraft that is found to be still operat- ing in 1967 could well affect our basic approach to sys- tem design philosophy. A notion of failure modes and mechanisms can also be obtained-again, to a restricted level-by discrete exercise of the various redundant pro- visions inherent in the design. In this connection, te- lemetry readings will provide invaluable diagnostic assistance. Redundancy exists for many of the critical elements required to make the 1967 tracking opportunity P-success. Except for the DC-15 condition, the redundant provisions of the spacecraft have not been utilized. The redundant elements assumed to be available are: 1. Cavity amplifier 2. Exciter B 3. Power frequency sources 4. Booster regulator 5. Analog-to-digital converter and pseudonoise generator 6. Gyro inertial roll control Limited determinations will also be possible regarding the following factors: 1. Solar cell performance trends 2. Clock accuracy 3. Deterioration of surfaces and finishes 4. Communication margin changes 5. Stability of stored picture data 81 JPL TECHNICAL REPORT NO. 32-881 Based on flight performance, there has been no reason for doubting that the spacecraft can continue operating for an extended period of time and, indeed, through 1967. Attitude control gas supply depletion is not antici- pated until late 1968; the low rate of gas consumption experienced so far should permit three-axis control of the spacecraft in 1967. And in order to receive telemetry data in 1967, it is not necessary to have command capability. b. Sampling of Interplanetary Conditions. Inherently, the greatest advantage lies in the basic scientific impor- tance of continued sampling of the interplanetary condi- tions for an extended period during a time of relatively high solar activity. Figures 54 and 55 show that the expected sampling will provide unique coverage across solar distances and activity greater than has been ex- perienced on any mission so far. 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I 9 u. u_ :.__ ;0 _-- .zz :zl'- _jz _mul .0 z JPL TECHNICAL REPORT NO. 32-881 93 JPL TECHNICAL REPORT NO. 32-881 SCHEDULE OF EVENTS ENCOUNTER ENCOUNTER- SEQUENCE PLANNING PROJECT REVIEW PROPOSED PLAN FINAL PLAN POST-ENCOUNTER PLANNING PROJECT REVIEW PROPOSALS BY DIVISIONS INITIATE REGULAR MEETINGS PROPOSED PLAN FINAL PLAN APR 5 12 19 26 5 I0 17 24. 51 7 14 21 28 5 12 19 26 2 /_ SCHEDULE • COMPLETION • • ACTIVITY SPAN SPAN PARTIALLY COMPLETE RESCHEDULED ENCOUNTER PLANNING SCHEDULE DATE:3-31-65/_gf /_ ! /._ PREPARED BY:_'_",,,_/<,L. REVISION I 6/10/65 94 JPL TECHNICAL REPORT NO. 32-881 95 JPL TECHNICAL REPORT NO. 32-881 96 L JPL TECHNICAL REPORT NO. 32-881 97 JPL TECHNICAL REPORT NO. 32-881 cJ iZ I "- II i_llml_llml_l qm i- I Ill .-.-.-. :,,,,-l lml)iill i % ) INtoi, umiii I I ) I II m i[ i _ ¢_-_ t t r i )i i i[ i -_ I iml i I iml i L _,,,-- _ mill i ill [ . 1 2o-t-FIIFFt-I_--'_i )ilia i IlIA I i _ -- - I i F i] [ i ii i _I i iNNA I lil J i . _ i I L I i k i i l l _,"..,Illi_ I im_ I . - ) liltl tIt) :o= Imlli lili l,,,._,.,,, llII)lllll =-i ImillIlill I •I IlilI Ilil,, ' )1[ "' """'"" I iN I I INL I I I _''I l lml I [il i I ) i i i i i_ i -,..._llmUl[[m[iI ¢ _ I iI I I )I I ",_ liEU PNil I _t IIll( Inl I ) --N_I I IN] I I iml I I .... _- _e_l ] iml i i iml I I - P -_i I lil I I lil I I I I [ I ) I I _ ] I.__.. ,,,-, ,,_,-, ,,,,- Iml Iml _.. I _'_,.,II ,.,_'_II -- Imliml ImlIml ; -_ Ill I i InN I I o_ __ > IIIllllF I[llllll I)1 iililil Ill ililli_ Q 98 JPL TECHNICAL REPORT NO. 32-881 99 JPL TECHNICAL REPORT NO. 32-881 APPENDIX C Summary of Reference Documents (Facsimile of a portion of an internal JPL document) 100 JPL TECHNICAL REPORT NO. 32-881 JPL Project Document No. 67 NASA DOCUMENTS Reports Title Booster Requirements Document, Revision 1 (Lewis Research Center} Mariner Mars 1964 Final Report, Volume I (U.S. Government Printing Office, Washington, D.C.) Mariner Mars 1964 Final Report, Volume II (U.S. Government Printing Office, Washington, D.C.) Mariner IV (Government Printing Office, Washing- ton, D.C.) Report from Mars No. Date BRD-4900 July 1964 Rev I In prep In prep NASA Facts 1965 Vol II No. 9 EP 39 In prep News Releases NASA to Begin Unmanned Mars Exploration This Year Mariner Mars 1964 Missions Scheduled for Early November Magnetometer Experiment Mariner IV Interim Science Findings, News Conference Solar Plasma Experiment Trapped Radiation Experiment Faint Radio Signals Provide Firm Lock on Mariner Mission Mariner IV Pre-Encounter Press Conference, Jet Propulsion Laboratory Mariner IV Press Conference 64-205 21 Aug 1964 64-266 29 Oct 1964 65-I17-B 19 Apr 1965 N65-21853 21 Apr 1965 65-I17A 19 Apr 1965 65-i17C 19 Apt 1965 65-198 16 Jun 1965 CSCL 17B N65-27472 22 Jun 1965 N65-30565 15 Jul 1965 CSCL 22A NASA News Releases available from NASA Scientific & Technical Information Div., Washington, D.C. - 2 - 101 JPL TECHNICAL REPORTNO. 32-881 JPL Project Document No. 67 News Releases (Cont' d) Title No. Date News Conference on Initial Scientific Interpretation N65-29786 29 Jul 1965 of Mariner IV Photography CSCL ZZA AFETR DOCUMENTS Air Force Systems Command Title No. Date Mariner-Mars 1964: Data Transmission Test Mariner-Mars 1964 Spacecraft: RF Systems Test Program Requirements Document Revision 3 Range Scheduling Handbook General Range Safety Plan Volume I General Range Safety Plan Volume II OD -4340 OD-4341 PRD-4300 Rev 3 AFM TCP 80-1 AFM TCP 80-2 Volume I AFM TCP 80-2 Volume II 18 Sep 1964 16 Sep 1964 15 Sep 1965 15 Jan 1963 5 Mar 1963 29 Oct 1962 Contact Section 293 for information about these documents. - 3 - 102 JPL TECHNICAL REPORT NO. 32-881 JPL Project Document No. 67 JPL TECHNICAL REPORTS Author Title No. Date Becker, R.A. Becker, R.A. Clarke, V.C., Jr., et al. Clarke, V.C., Jr. Coyle, G. Dawson, K.M., et al. Dallas, S.S. Easterling, M.F. Easterling, M. Goldstein, R. Groudle, T. Schmitz, D. Hunter, J.H. James, J.N., et al. Analysis of Solar Panel Effect TR 32-687 I Jun 1965 Design and Test Performance of Mariner IV TelevisionOptical System TR 32-773 1 Jul 1965 Design Parameters for Ballis- tics InterplanetaryTrajec- tories, Part l: One Way Transfers to Mars and Venus TR 32-77 16 Jan 1963 Trajectory Design for Ranger Miss ions TR 32-471 1 Mar 1965 Mariner IV Science Platform Structure and Actuator Design, Development and Flight Per- formance TR 32-832 15 Nov 1965 Reliability Considerations in the Design, Assembly, and Testing of the Mariner IV Power System TR 32-729 l Jul 1965 High Energy Earth to Mars and Return Trajectories TR 32-803 in prep A Long Range Precision Ranging System TR 32-80 t0 Jul 1961 The Effect of the Interplanetary TR 32-825 Medium on S-Band Telecommu- nications Development of the Mariner C Post-Injection Propulsion System Mariner Mars 1964 Telecom- munication System 1Sep 1965 TR 32-830 In prep Mariner IV Mission to Mars, Part I TR 32-836 1 Dec 1965 TR 32-782 15 Sep 1965 -5 - 103 JPL TECHNICAL REPORT NO. 32-881 JPL Project Document No. 67 JPL TECHNICAL REPORTS (Cont' d) Author Title No. Date Kliore, A., et al. Mathison, R.P. Pickering, W.H., etal. Schiffer, R. A, Schutz, F.L., et al. Thostesen, T.O. Lewis, D.W. Wong, R.Y. Project Team Determination of Some Physical Properties of the Atmosphere of Mars from the Changes in Dop- pler Signal of a Spacecraft on an Earth Occultation Trajectory Mariner Mars 1964 Telemetry and Command System The Mariner IV Mission to Mars Correlation of Launch Vehicle Wind Tunnel Aerodynamic Noise with Flight Vibration Data, Revision 1 The Mariner Mars Science Subsystem The Mariner Mars 1964 Absorptivity Standard A Fiber Optical System for Space Application Mariner Mars 1964 Project Report, Mission and Spacecraft Development; Volume I: Basic Report Mariner Mars 1964 Project Report, Mission and Spacecraft Development; Volume II: Appendixes Mariner Mars 1964 Project Report, Mission Operations Mariner Mars 1964 Project Report, Spacecraft Perfor- mance and Analysis TR 32-674 16 Oct 1964 TR 32-684 i Jun 1965 TR 32-782 15 Dec 1965 Part II TR 32-619 15 Sep 1964 Rev I TR 32-813 15 Feb 1966 TR 32-734 1 Mar 1966 TR 32-646 31 Aug 1964 TR 32-740 In prep Vol I TR 32-740 In prep Vol II TR 3Z-881 In prep TR 32-882 In prep 6 - 104 JPL. TECHNICAL.REPORT NO. 32-881 JPL Project Document No. 67 JPL TECHNICAL REPORTS (Cont' d) Autho r Title No. Date Project Team Mariner Mars 1964 Project Mariner Mars 1964 Project Report, Television Experiment, Part i: investigators' Report Mariner Mars 1964 Project Report, Television Experiment, Part II: Picture Element Matrices TR 32-883 In prep TR 32-884 In prep Part I TR 32-884 In prep Part II 7 105 JPL TECHNICAL REPORT NO. 32-881 JPL Project Document No. 67 JPL TECHNICAL MEMORANDA Author Title No. Date Batchilder, R.R. Wada, B.K. Haynes, N.R. Gordon, H.J. Bastow, J.G., Johnson, N.E. Pedersen, E.S. Pefley, R.K. Wilson, J.H. Renzetti, N.A. Renzetti, N.A. Jr., et al. Stiffness Matrix Structural Analysis A Study of the Probability of Depositing Viable Organisms on Mars During the Mariner ' 64 Mission Proceedings of the Magnetic Workshop, March 30 to April 1, 1965, CIT, Pasadena Investigation of Fiberglass Shroud Materials Heat-Sterilizable Power Source Study for Advanced Mariner Missions Temperature Control, A Case History of the Mariner Space- c raft The Odyssey of Mariner IV Tracking and Data Acquisition Support for Mariner Mars 1964 Volume I. Near Earth Trajec- tory Phase Tracking and Data Acquisition Support for Mariner Mars 1964, Volume II. Cruise to Post- Encounter Phase TM 33-75 12 Feb 1962 TM33-194 23 Oct 1964 TM 33-216 15 Sep 1965 TM 33-214 I Apr 1965 TM 33-180 I Jul 1964 TM 33-189 i Mar 1965 TM 33-ZZ9 Jul 1965 TM 33-239 I Jan 1965 Voll TM 33-Z39 In prep Vol II -8 - 106 JPL TECHNICAL REPORT NO. 32-881 JPL Project Document No. 67 OTHER JPL FORMAL REPORTS IN PREPARTkTION Antenna Structure, Design and Development Basic Structure, Design and Development Design and Manufacturing of Mariner C Solar Panels Design, Mechanization, and Flight Analysis of the Mariner C Power System Design of Mariner Power Operational Support Equipment Design of the Mariner C Power Conversion Equipment Development Spacecraft Utilization Effects of Dust Particles on the Canopus Tracker Electronic Cabling Electronic Packaging Filtering Techniques for Noise Suppression in Quasi-Balanced Circuits Manufacturing and Quality Assurance of the Mariner C Conversion Equipment Mariner C Battery, Final Report Mariner C Environmental Facilities and Test Program Summary Mariner C Post-lnjection Propulsion System R_galator Development Mariner C Post-Injection Propulsion System Squib Development Mariner C Post-Injection Propulsion System Squib/Valve Development MC-4 Timing and Sequencing Subsystem Mariner IV Flight Path and Its Determination from Tracking Data Mariner Mars Temperature Control Subsystem Mechanical Configuration Performance Evaluation of Mariner C Solar Panels Propulsion System Components Solar Panel: Structure, Boost Restraint, and Deployment Spacecraft Inertial Property Determination -9- 107 JPL TECHNICAL REPORT NO. 32-881 JPL Project Document No. 67 OTHER JPL FORMAL REPORTS IN PREPARATION (Cont' d) Temperature Control Design, Testing, and Flight Results Temperature Control Hardware Design and Development Total Propulsion System Configuration 10 - 108 JPL TECHNICAL REPORT NO. 32-881 JPL Project Document No. 67 SPACE PROGRAMS SUMMARY, VOLUMES III The Deep Space Network No. Page Title Date 37-29 38-39 30 Sep 1964 37-30 34-35 30 Nov 1964 37-30 36 30 Nov 1964 37-31 32 31 Jan 1965 37-31 33-34 31 Jan 1965 37-32 9-10 31 Mar !965 37-33 32-38 31 May 1965 37-25 i8-i9 M,irl,ei C I00 kw Transmitter 31 Jan 1964 37-26 22 Mariner C 100-kw Transmitter 31 Mar 1964 37-27 37-38 SFOF Operations 31 May 1964 37-27 52 Mariner C Transmitter Development 31 May 1964 37-28 43-45 Mariner C Transmitter Development 31 Jul 1964 37-29 38 Ground Instrumentation for Mariner C Occultation 30 Sep 1964 Measurements Mariner C Transmitter Development Ground Instrumentation for Mariner C Mariner C 100-kw Transmitter Ground Instrumentation for Mariner C Occultation Experiment Mariner C 100-kw Transmitter Ground Instrumentation for Mariner iV Occultation Experiment Ground Instrumentation for Mariner IV Occultation Experiment 37-33 56-58 Venus Station Mariner IV Support 31 May 1965 37-33 58-62 Venus Station Mariner IV Encounter Receiver 31 May 1965 37-34 41-44 Ground Instrumentation for Mariner C Occultation 31 Jul 1965 Experiment Mariner IV Operations Mariner Master Data Library Ground Instrumentation for Mariner IV Occultation Experiment Venus Station Mariner IV Encounter Receiver and X-Band Lunar Radar 37-34 3-4 31Jul 1965 37-35 34-37 30 Sep 1965 37-35 39-42 30 Sep 1965 37-35 42-45 30 Sep 1965 -11 109 JPL TECHNICAL REPORT NO. 32-881 I a JPL Project Document No. 67 SPACE PROGRAMS SUMMARY, VOLUMES IV Supporting Research and Advanced Development No. Page Title Date 37-25 49 Mariner Configuration Study 29 Feb 1964 37-26 74-75 Welded Electronic Packaging 30 Apr 1964 37-27 92-97 Carrier Collection and Spectral Response of 30 Jun 1964 Radiation Detectors 37-28 101-108 Planet Acquisition-Range Prediction of a Planetary 31 Aug 1964 Scan System 37-30 131 High- and Low-Field Characteristics of the Helium 31 Dec 1964 Magnetometer 37-33 84-90 Printed Conductor Assembly Substitution for 30 Jun 1965 Mariner C Upper Ring Harness SPACE PROGRAMS SUMMARY, VOLUMES VI Space Exploration Programs and Space Sciences 37-17 29-31 Mariner Command Subsystem 31 Oct 1962 37-18 70-72 Planetary Scan Systems 31 Dec 1962 37-19 43-46 Mariner Mars 1964 Design 28 Feb 1963 37-19 81-82 Mariner Mars 1964 Logically Controlled Shutter- 28 Feb 1963 and-Filter Wheel 37-20 27-34 Mariner Mars 1964 Design 30 Apr 1963 37-20 I04-i08 Mariner Mars 1964 Photographic Subsystem 30 Apr 1963 37-20 i12-I13 Mariner Mars 1964 Television Subsystem Optical 30 Apt 1963 Subassembly 37-21 32 Launch and Arrival Date Considerations 30 dun 1963 37-21 32-42 Spacecraft Design 30 Jun 1963 37-21 75-78 Planetary Scan Subsystem for Mariner Mars 1964 30 Jun 1963 37-22 17-18 Spacecraft Mission Simulation 31 Aug 1963 -12- 110 JPL TECHNICAL REPORT NO. 32-881 JPL Project Document No. 67 SPACE PROGRAMS SUMMARY, VOLUMES VI (Cont' d) No. Page Title Date 37-22 18 Mariner Mars 1964 Systems Test Plans 31 Aug 1963 37-22 19-24 Mariner Mars 1964 Spacecraft Deslgn 31 Aug !963 37-22 53-56 TV Camera Shutter 31 Aug 1963 37-23 17-18 Mariner Mars 1964 TV System Test 31 Oct 1963 37-23 18-26 Mariner Mars 1964 Spacecraft Design 31 Oct 1963 37-23 42-49 Operational Support Equipment 31 Oct 1963 37-24 14-22 Spacecraft Design and Development 31 Dec 1963 37-24 71-87 Mariner C Scientific Instrumentation 31 Dec 1963 37-25 12-13 Systems Testing 31 Jan 1964 37-25 13-14 Atlantic Missile Range Launch Complex Facility 31 Jan 1964 Design 37-25 14-18 Spacecraft Design and Development 31 Jan 1964 37-25 41-43 Mariner Television Subsystem Optical Subassembly 31 Jan 1964 37-25 43-46 Mariner C TV Development Support _quipment 31 Jan 1964 37-25 53-56 Mechanization of the iNonreal Time Data Auto- 31 Jan 1964 mation System 37-25 56-58 Planetary Simulation for Mariner C Planetary 31 Jan 1964 Scan Subsystem 37-26 20 Flight Analysis Program 31 Mar 1964 37-26 20-21 Systems Testing 31 Mar 1964 37-26 21-23 Spacecraft Development and Testing 31 Mar 1964 37-26 55-69 Mariner C Magnetometer 31 Mar 1964 37-26 70-72 Improved Slow-Scan Vidicon Camera Tubes 31 Mar 1964 (Mariner C) 37-26 72-28 Effects of Bit Errors on Pulse Code Modulated 31 Mar 1964 (PCM) Television Picture Data (Mariner C) -13 - 111 JPL TECHNICAL REPORT NO. 32-881 JPL Project Document No. 67 SPACE PROGRAMS SUMMARY, VOLUMES VI (Cont' d) No. Page Title Date 37-28 48-51 31Jul 1964 37-29 34-37 30 Sep 1964 37-29 37-44 30 Sep 1964 37-29 59-63 30 Sep 1964 37-30 18-20 30 Nov 1964 37-30 20 30 Nov 1964 37-30 20-21 30 Nov 1964 37-31 16-18 31 Jan 1965 37-31 18-19 31 Jan 1965 37-31 19-21 31 Jan 1965 37-3! 54-55 31 Jan 1965 37-27 10-11 System Testing 31 May 1964 37-27 If-15 Spacecraft Design and Development 31 May 1964 37-27 15-17 Three-Axis Spacecraft Simulator 31 May 1964 37-27 41-44 Planetary Scan Subsystem (Mariner C) 31 May 1964 37-28 18-21 Spacecraft Systems Testing 31 Jul 1964 37-28 21-29 Design and Development 31 Jul 1964 37-28 43-48 Mariner TV Subsystem Science and Engineering 31 Jul 1964 Calibration The Mariner Vidicon as a Ruggedized Space Component Spacecraft Systems Testing Design and Development Mariner C TV Subsystem Camera-Shutter Electronics Spacecraft Systems Testing Spacecraft Thermal Testing Design and Development Mission Operations Power Subsystem Operation During the Mariner Mission Developmental and Testing Activities Magnetic Shields to Support Mariner C Magneto- meter Testing 37-32 14 Mariner IV Mission Operations 31 Mar 1965 37-32 14 Mariner IV Power Subsystem Performance 31 Mar 1965 37-32 15-16 IAfe Testing Activities 31 Mar 1965 14 - 112 JPL TECHNICAL REPORT NO. 32-881 JPL Project Document No. 67 SPACE PROGRAMS SUMMARY, VOLUMES VI (Cont' d) No. Page Title Date 37-32 37-41 Mariner C TV Field Support Operations 31 Mar 1965 37-32 41-43 Mariner C Magnetometer: Development of 31 Mar 1965 Modified Helium Lamps and Cells 37-32 44-48 Methods of Determining the Characteristics and 31 Mar 1965 Performance of Radiation Detectors Used in the Mariner C Planetary Scan Subsystem 37-32 49-55 Advanced Mariner Data Automation System 31 Mar 1965 Development 37-33 18 Mariner IV Space Flight Operations 31 May 1965 37-33 18-22 Mariner IV Attitude Control Subsystem 31 May 1965 Pe rfo rmance 37-33 22 Development of a Star Identification Procedure 31 May 1965 37-34 20-21 Mariner IV Space Flight Operations 31 Jul 1965 37-34 21-22 Mariner IV Radio Subsystem Performance 31 Jul 1965 37-34 27-39 Mariner IV Ionization Chamber Failure Analysis 31 Jul 1965 37-34 39-41 N1ariner IV Plasma Probe Jeallure Analysis 31 Jul i_u'_5 37-34 41-44 Mariner IV TV Interpretation Studies 31 Jul 1965 37-34 44-47 Spurious TV Image Phenomena 31 Jul 1965 37-34 48-49 Space Sciences Encounter Planning Activities 31 Jul 1965 37-35 10 Mariner IV Space Flight Operations 30 Sep 1965 37-35 I0-12 Mariner IV Spacecraft Performance During The 30 Sep 1965 Mission Future Operations for Mariner IV Mariner Master Data Library Total Radiation Dose Experienced by the Mariner IV Spacecraft as Measured by the Ion Chanlber Experiment 37-35 12 30 Sep 1965 37-35 22 30 Sep 1965 37-35 35-37 30 Sep 1965 -15 113 JPL TECHNICAL REPORT NO. 32-881 JPL Project Document No. 67 SPACE PROGRAMS SUMMARY, VOLUMES VI (Cont'd) No. Page Title Date 37-35 37 -44 Mariner IV Magnetic Measurements Inside the Earth's Magnetosphere, Magnetic Tail, and Magnetosheath 30 Sep 1965 16 - 114 JPL TECHNICAL REPORT NO. 32-881 JPL Project Document No. 67 OPEN LITERATURE Author Title Publication Date Alexander, W.M., et al. Anderson, H.R. Billingsley, F.C. Bouvier, H.K. Casani, J.R., et al. Christensen, E. Coughey, T.K. Dawson, K.M., et al. Drummond, A.J. Laue, E.G. Zodiacal Dust: Measure- ments by Mariner IV Spacecraft Description and Encounter Sequence Digital Video Processing at JPL (10th Annual Symposium, San Francisco, Calif., 16-20 August 1965) The Mariner IV Attitude Control System Mariner IV: A Point of Departure Ranger and Mariner Tem- perature Control Experi- ences (AIAA Thermophysics Specialists Conference Sponsored by NBS and AF Materials Lab, Monterey, California, 13-15 Sep 1965) Classical Normal Modes in Ramped Linear Dynamic Systems Design and Reliability Con- siderations for the Mariner Mars 1964 Spacecraft Power System (AIAA Unmanned Spacecraft Meeting, Los Angeles, I-4 March, 1965) The Extraterrestrial Solar Spectrum and its Simulation in Space Environmental Test System (IES Solar Simulation Symposium, Los Angeles, January 1964 Science 149:1240-1241 Science 149:1226-1228 10 Sep 1965 10 Sep 1965 Conference To be pub Proceedings Space Res & Eng (France) I:7-8:41 Astronaut & Aeronaut 3:8:16-24 1965 3 Aug 1965 Conference To be pub Proceedings J Appl Mech Jun 1960 269-271 AIAA Pub CP-12, 323-332 1965 Conference Oct 1964 Proceedings -18 - 115 JPL TECHNICAL REPORT NO. 32-881 Q JPL Project Document No. 67 OPEN LITERATURE (Cont' d) Author Title Publication Date Ewing, A. Fawcctt, W.G., et al. Fink, D.E. Forney, R.G., et al. Foss, K.A. Gianopulos, G.N. Goldstein, R.M. Haynes, N.R., et al. Heller, G.B. Herl-nan, N.H. Lingon, U.S. Highstrete, B,A. Hyde, J.R. Canals May Be Craters Scientific Exploration with Mariner IV Mars Vehicle Becomes Major Scientific Program Mariner IV Maneuver and Control Coordinates which Uncouple the Equations of Motion of Damped Linear Dynamic Systems Use of the Computer in Ranger and Mariner Projects (IBM Symposium on Com- puter Aid Experimentation, Poughkeepsie, New, York, 11-13 October 1965) JPL Radar Observations of Mars Mariner IV Flight Path to Mars A First for Therrno Physics (Review) Mariner IV Tinling and Sequencing Ceranlic Traveling Wave- tubes for Space Vehicles Mariner Mars 1964 Acous- tically Induced Vibration Environnqent (35th Symposiunl o,_ Shock Vibration, Shock and Vibration Inforn_ation Center, New Orleans, 25-28 Oct 1965) Sci. NL 88:103 14 Aug 1965 Astronaut & Aeronaut 3:10:22-28 Oct 1965 Aviation W. 82:116-18+ 15 Mar 1965 Astronaut & Aeronaut 3:10:36 Oct 1965 J Appl Mech Sep 1958 361 -364 Conference To be pub Proceedings by IBM Sc ienc e 150:3704:1715-1717 Astronaut & Aeronaut 3:28-33 24 Dec 1965 Jun 1965 Astronaut & Jan 1966 Aeronaut p 42 Astronaut & Aeronaut 3:10:40 Oct 1965 Microwave J Jan 1964 7:49-57 Conference To bc pub Proceedings -19- 116 JPL TECHNICAL REPORT NO. 32-881 JPL Project Document No. 67 OPEN LITERATURE (Cont' d) Author Title Public ation Date Hyde, J.R. Systems Testing (1965 J Environ To be pub Educationai Seminar, " .... _^" Research Center, Dec 1965) James, J.N. Kliore, A., et al. Kliore, A., et al. Kliore, A. 3. Lear, J. Leighton, R. B. , et al° Lewis, Donald W. Lindsey, R. Malling, L: R. Mailing, L.R. Managing the Mariner Mars Project Mariner IV Occultation Experiment Occultation Experiment: Results of the First Direct Measurement of Mars Atmo- sphere and Ionosphere Preliminary Results of the Mariner IV Occultation Measurement of theAtmo- sphere of Mars (GIT/JPL Lunar and Planetary Gon- ference, 13-18 Sep 1965) Mariner IV' s Expense Account: Increase Due to Extreme Tenuity of Martian Atmosphere Mariner IV Photography of Mars: Initial Results Maintaining Thermal Control in Mariner IV Mars Atmosphere Probe Proposed Slow Scan Vidicon as an Interplanetary Imaging Device Space Astronomy and the Slow Scan Vidicon Astronaut & Aeronaut 3:8:34-41 Astronaut & Aeronaut 3:72-80 Science 149: 1243- 1248 Conference Proceedings Sat R 48:35 Science 149:627-30 Astronaut & Aeronaut 3:10:30 Miss _: Roc Advances in Electronics and Electron Physics (Academic Press) J SMPT Aug1965 Jul 1965 10 Sep 1965 1965 7 Aug1965 6 Aug1965 Oct 1965 8 Feb1965 To be pub Nov1963 -20 - !17 JPL TECHNICAL REPORT NO. 32-881 JPL Project Document No. 67 OPEN LITERATURE (Cont'd) Author Title Public ation Date Mathison, R.P. Meisenholder, G.W. Miller, B. Normyle, W.J. O' Gallagher, et al. Pay, R. Pickering, W.H. Schutz, F.L. Schmuecker, J.D. Wilson, J. N. Schneiderman, D., et al. Shipley, W.S. Maclay, J.E. Mariner Mars 1964 Telemetry and Command System The Mariner IV Canopus Tracker (XVI International Astronautical Congress, Athens, Greece, 13-18 September 1965) JPL Facing Marine_ C Avionics Problems Planetary Exploration Hopes Buoyed by Mariner Flight Search for Trapped Electrons and a Magnetic Moment at Mars by Mariner IV VAD Group Processed Mariner Photos Mariner IV' s Flight to Mars Integration of Scientific Instruments into Mariner IV (XVI Int' 1 Astronautical Congress, Athens, Greece, 13-18 September 1965) Structural and Mechanical Design of Mariner Mars Recent Mariner Spacecraft Missions (AIAA Unmanned Spacecraft Meeting Los Angeles, 1-4 March 1965) Mariner IV Environmental Testing IEEE Spectrum Jul 1965 2:76 -84 Conference To be pub Proceedings (Poland) 1966 Aviation W 29 Jun 1965 80:16-17 Aviation W 9 Aug 1965 83:86 -7 Science i0 Sep 1965 149:1233-1239 Miss & Roc 9 Aug 1965 17:28+ Astronaut Oct 1965 Aeronaut 3:10:20 Conference To be pub Proceedings (Poland) 1966 Astronaut & Aeronaut 3:8:26-33 AIAA Pub CP-12, 385-394 Astronaut Acta II Astronaut & Aeronaut 3:8:42-48 Aug 1965 1965 Aug 1965 21 - 118 ,o | JPL TECHNICAL REPORT NO. 32-881 JPL Project Document No. 67 OPEN LITERATURE (Cont' d) Author Title Publication Date Scull, J.R. Guidance and Control of the Conference To be pub Mariner Planetary Space- Proceedings craft (IFAC Symposium on Peaceful Uses of Automatic Controls in Space, Stavanger, Norway, 2-Z4 June 1964) Sjogren: W.J. Astrodynamical Results From J Spacecraft Sep- Trask.. D.W. the Radio Tracking of the Rockets 2:689-697 Oct 1965 Mariner Venus and Ranger Lunar Missions Sloan, R.K., et al. Mariner IV Flight Path to Astronaut& Jun 1965 Mars Aeronaut Smith, E.J., et al. Magnetic Field Measure- Science I0 Sep 1965 ments Near Mars 149:1241-1242 Smith, E.J. Smith, E.J. Soffen, G.A. Stone, L. Stone, L. A Review of Lunar and Planetary Magnetic Field Measurements Using Space Probes (NATO Advanced Study Institute Meeting on Planetary and Stellar Magnetism Magnetism and the To be pub Cosmos, Conference Proceedings (Oliver & Boyd, Edinburgh, Scotland) Measurement of Magnetic Fields in the Vicinity of Mag- netosphere and in Interplan- etary Space Prelim. Results from Mariner IV (Sixth COSPAR Meeting) Space Kesearch VI To be pub 1966 Mars Biological Mission: Possibilities and Constraints (AAS Symposium on Unman- ned Exploration of Space, Denver, Colo., 8-10 Feb- ruary 1965) Bioscience To be pub Mariner Data May Limit Voyager Payload Aviation W g Aug 1965 83:55- Mariner Design Modified by Mars Flyby Aviation W 6 May 1965 78:50-1+ -22 - 119 JPL TECHNICAL REPORT NO. 32-881 JPL Project Document No. 67 OPEN LITERATURE (Cont' d) Author Title Public ation Date Taylor, H. Webb Says Mariner Winning Miss & Roc 15 Febl965 Mars Race: Zond a Month 16:15 Behind Thomas, G.M. Experimental Measurements AIAA CP-9 Menard, W.A. of Nonequilibrium andEquili- brium Radiation from Plan- etary Atmospheres (AIAA Entry Technology Conference, Williamsburg and Hampton, Va., 12-14 October 1964) Van Allen, J.A. , et al. Wainwright, L. Watkins, H.D. Absence of Martian Radiation Science Belts and Implications Thereof 149:1228-1233 Our Encounter with Mars: Successful Combination of Men and Machines Life 59:14 JPL Debating Alternate Methods for Contacting Mariner IV in 1967 Aviation W 83:32 #atts, R.N., Jr. Mariner Flight Continues Sky & Tel 29:95 -6 Watts, R.N., J'r. Mariner IV Completes Mars Sky & Tel Mission 30:136 -8 Watts, R.N., Jr. Mars Observations Wanted: Sky & Tel Mariner IV Probe 29:150 Watts, R.N., Jr. Results from Mariner Sky & Tel 26:195t Way, John, Sr. Analog-Digital Conversion Space/Aeron of TV Data on Mariner IV 44:83-88 Welnick, R.A. Wright, F.H. Wilson, J. ti. Assuring Quality and Reli- ability for Mariner IV Astronaut& Aeronaut 3:8:50 Space Probes Encyc Sci & Technology 1966 Yearbook (McGraw-Hill) 1965 I0 Sep 1965 6 Aug1965 2 Aug 1965 Feb1965 Sep 1965 Marl965 Oct 1963 Aug 1965 Aug 1965 1966 -23 - 120 JPL TECHNICAL REPORT NO. 32-881 JPL Project Document No. 67 OPEN LITERATURE (Cont' d) Author Title Public ation Date Younkin, R.L. A Search for Limonite Near In/rated Spectral Features on Mars Astrophys J May 1966 Craters Found on Mars Journey to July: Mariner IV Mariner Photos to Be Studied with Eye to Voyager Mariner IV Measurements Near Mars: Initial Results Mariner IV Photographs of Mars Mariner IV Sensors Relay Excellent Data Mars in Focus: Mariner IV Probe Martian Surface Shows Moon- Like Quality Moonfac ed Mars Moon-faced Mars: Concerning Pictures Taken by Spaceship Mariner IV Notes and Comment: News from Mariner IV Rendezvous with Mars Section of Planetary Sciences, Symposium on Mariner IV Interplanetary Results Symposium on Mariner IV Re su Its Sci NL 88:82 Newsweek 64:67 Miss & Roc 17:16-17 Science 149:1179, 1226-48 Sky & Tel 30:155-61 Aviation W 81:26-7 Sr Schol 87:19-20 Aviation W 83:30-1 Newsweek 66:58 Time 86:58 New Yorker 41:17 Sci Digest 57:24-7 Am Geophys Union Trans 46:113 Am Geophys Union Trans 46:532 7 Aug 1965 7 Dec 1965 9 Aug 1965 I0 Sep 1965 Sep 1965 7 Dec 1964 16 Sep 1965 2 Aug 1965 9 Aug 1965 6 Aug 1965 7 Aug 1965 Jan 1965 Mar 1965 Sep 1965 - 24 - 121 JPL TECHNICAL REPORTNO. 32-881 JPL Project Document No. 67 OPEN LITERATURE (Cont'd) Author Title Public ation Date Terrain of Neighbor Mars: Pictures Taken 135 Million Miles Away Life 59:62A-62C What Scientists Hope to Learn U.S. News from Mars Shot: Mariner IV 57:18-19 With an Eye on Canopus: U.S. Newsweek 64:58 Mariner IV 6 Aug 1965 14 Dec 1964 14 Dec 1964 _5 - 122 t6 JPL TECHNICAL REPORT NO, 32-881 JPL Project Document No. 67 CONFERENCE PAPERS Author Title Occasion _nderson, H.R. Anderson, H.R. Anderson, H.R. Anderson, H.R. Bell, D.E. Billingsley, F.C. Cherry, W.R. Zoutendyke, J.A. Conrad, A.G. Dow, P.C., Jr. Eckman, P.K., et al. Interplanetary Radiation Measured by Mariner II and'IV Preliminary Results of Mariner IV Ion Chamber Experiment Results from the Mariner IV Ion Chamber Experiment The Space Environment Measured by Mariner IV (Summary of Mate- rial presented at AGU Meeting, 19-20 April 1965 Mariner IV TV Camera and Mission Results Mariner Mars 1964 Television The State of the Art in Solar Cell Arrays for Space Electrical Power Design and Flight Operation of Mariner IV Exploration of Interplanetary Space by Unmanned Vehicles Some Possible Means of Acquir- ing Increased Information Return from Mars Seminar, Rice University Cosmic Ray Conference January 1965 Intl. Conf. on Cosmic Rays, London, England 6-17 Sep 1965 AIAA International Symposium on Space Electronics, Miami Beach, Florida, 2-5 Nov 1965 Data Reduction 6, Com- puter Working Group, of Inter-Range Instrumenta- tion, Los Angeles, 20 Oct 1965 AIAA 64-738, Sep 1964 AIAA 2nd Annual Meeting San Francisco, 26-29 July 1965 Fifth European Space FlightSymposium, Munich, West Germany 19-22 July 1965 Symposium on Unmanned Exploration of the Solar System 8 Feb 1965 26 - 123 JPL TECHNICAL REPORT NO. 32-881 'r JPL Project Document No. 67 CONFERENCE PAPERS (Cont' d) Author Title Oc c a sion Gabron, F., et al. Goble, M.E. Earle, W.G. Gordon, H.J. Michel, J.R. Happe, R.A. Herriman, A. LaBaw, C. C. Hyde, J.R. Hyde, J.R. Hyde, J.R. Jodele, J. Lewis, D.W. Thostesen, T.O. Thermal Scale Modeling of a Modified Prototype of the Mariner Mars 1964 Spacecraft System Test Program for the Mariner to Mars Spacecraft Midcourse Guidance for Mariner Mars 1964 Materials Engineering for Inter- planetary Probes The Mariner IV Photographic Mission Managing the Mariner Mars Spacecraft Test Program Mariner IV Mission to Mars The Mariner IV Mission to Mars Packaging Techniques for Mariner IV Spacecraft Components Mariner-Mars Absorptanc e Experiment Paper 65-386 AIAA Znd Annual Meeting, San Francisco, 26-29 Jull965 AIAA-AAS Conference: Stepping Stones to Mars, Baltimore, Maryland, 28 Mar 1966 Paper 65-402 AIAA 2nd Annual Meeting, San Francisco, 26-29Jui 1965 1965 Metals/Materials Congress, Detroit, 22 Oct 1965 98th Conference of Motion Picture and TV Engineers, Montreal, Canada, I Nov 1965 Richards-Gebaur AFB Management Group, Kansas City, 27 Jan 1966 Armed Forces Communications Elec- tronics Assn, Kansas City, 27 Jan 1966 18th Annual Mid-An, erica Electronics Conference, Kansas City, Mo., l8 Nov 1965 Intl. Conf. on Components &MaterialsUsed in Elec- tronics Engineering, London, 17-25 May 1965 Paper 65-650 AIAA Conference on Thernlal Physics, Monterey, Calif. , 13-15 Sep 1965 27 - 124 JPL TECHNICAL REPORT NO. 32-881 JPL Project Document No. 67 CONFERENCE PAPERS (Cont' d) Author Title Oc c as ion Maclay, J.E., et al. Mailing, L.R. Vacuum Capability Requirements for the Mariner T_sting Program Data Compression for Planetary Photography Mailing, L.R. Time Bandwidth Compression for Planetary Pictures Meghreblian, R.V., Science Payloads from Mariner II et al. to Mariner IV Meisenholder, G.W. Stabilization of Ranger and Mariner Reconnaissance Systems Porter, R.W. Space Research in the United States, 1964 Olbeter, A., et al. A Sensitive S-Band Noise Meter Developed for the Mariner Mars ' 64 Spacecraft Program Schiffer, R.A. Hyde, J.R. Schneiderman, D. The Mariner Mars 1964 Spacecraft Acoustically Induced Vibration Environment Mariner IV - The Design, Applica- tion and Results Schneiderman, D. Sloan, R.K. Review of Mariner 1962 and 1964 Spacecraft Design and Operational Planning The Interplanetary Mission of Mariner IV ASTM 5th Pacific Nat Meeting, 3 Nov 1965 Paper 97-8 SMPTE Conference, Los Angeles, 29 Mar - 2 Apr 1965 SPIE Technical Sympo- sium, San Francisco, 28 Aug 1964 NASA-AFCRL Conference on Space Science and Technology, Roanoke, Va., 23-27 Aug 1965 SPSE Conference, Cleveland, Ohio 17 May 1965 COSPAR Plenary Meeting, Buenos Aires, Argentina, 11-12 and 19-21 May 1965 IEEE 7th Annual Sympo- Compatability, New York City, 28 -30 June 1965 5th International Congress on Acoustics, Liege, Belgium, 7-14 Sep 1965 6th Intl. Symposium on Space Technology Science, Tokyo, Japan, 29 Nov - 7 Dec 1965 Conference on Mars and Venus, Virginia Poly- technic Institute, Sep1965 AGU Meeting, 19 April 1965, Washington, D.C. - Z8 - t25 JPL TECHNICAL REPORT NO. 32-881 ¢ JPL Project Document No. 67 CONFERENCE PAPERS (Cont' d) Author Title Occasion Snyder, C.W. Soffen, G.A. South, J.F. Thomas, G.M. Tito, D.A. Wiggins, C.P. Younkin, R. h. Some Aspects of the Mariner IV Solar-Plasma Observations Life Detection Experiments Space-Simulated Testing of Mariner Spacecraft Autopilot Systems With Rocket Motor Firing Ultraviolet Radiation Program at JPL Trajectory Design for the Mariner 1964 Mission High Power Microwave Trans- mitters for Deep Space Communica-- tion with Unmanned Spacecraft Spectral Energy Distribution and Reflectivities of Martian Maria and Continents in the Visible and Near Infrared AGU Meeting, 19 April 1965, Washington, D.C. Lecture in UCLA course of study (four campuses): Horizons in Space Sciences, Spring 1964 (Series of lectures to be pub by UC, Berkeley) Paper 63-343 AIAA Guidance and Control Conference, MIT, Cambridge, Mass., 12-14 Aug 1963 NASA Research Div, Office of Advanced Research & Technology, Fluid Physics Contractors Conference on Ultraviolet Radiation and Heat Trans- fer, 31 Aug 1965 Paper 65-516 AIAA 2nd Annual Meeting, San Francisco, 26-29 July 1965 1966 National Telemetering Conference, Boston, Mass., 10-12 May 1966 (Proceedings to be pub- lished) AAS Meeting, Ann Arbor, Mich., Aug 1965 (Abstract in Astronom J Nov 1.965) - 29 - 126 JPL TECHNICAL REPORT NO. 32-881 APPENDIX D Encounter Sequence of Events 127 JPL TECHNICAL REPORT NO. 32-881 cq b3 0 b-1 u_ rj _D Z o_ D 0 Z _ 0 Z 0 o .,-t ..H ¢) 'O -,-.I ,-4 0 4o °r.I % .el .,-t 4J 0 ._ 0 0 0 ..¢..t ,-4 0 & 0 t_ r_ r_ ,H _d _ tO o _ ._ _ ,_14-1 4-' ¢J •,a _ f_ u) -,-4 ,-_ _ 0 _ o_ o o_o o _ _ .. ® 0 0 0 0 _ r_ _C/.1_ _ ._ -_ o • "13 0 1 •- -- _ _ _ _ _, _ _ _" _ 0 0 ® 0 •_ _ 0 0 o _D 0 _ • _., ._ .,_ _ 0 _ ,-_ 0 c_ ._ _ _ .._ 0 0 0 r,, t:_ 0_ v rj C_ _._ _ 0 -.-4 0 _ o 0 C_ _ _ _ ._ o 0 ,.-q 0 ;3 0 0,_ _d _ c_ 03 0_ ,-r, 0 r_ ° r.j _ o o _ _ _ o _ _ _ 0 0_ II Or., o _ °,..t H _ 0 0 0 0 0 _ r-.4 ,--I ,--I _ g o 0 0 .o "_ 4o o o e, % ! ! ! 128 JPL TECHNICAL REPORT NO. 32-881 000000000_ I I 6 Z_p 0 Z_ -o_o----p_o O_J VVVVVW=_L-_-L. 1¢ 8 :+=..._ ++ +- - _J, o l,,, IiI i,,i ! 129 JPL TECHNICAL REPORT NO. 32-881 ,¢,> w t_ W m o- I¢ , ,_ _ o. Z ! 130 JPL TECHNICAL REPORT NO. 32-881 We; N • .,J IL W oo oo oo oo oo 3_g_88 5 O _-_ _ o °°°''" ===¢¢ 5 5" _= g g °° °ooooo °oo°o o oo o o • | a • • • Id L ! 131 JPL TECHNICAL REPORT NO. 32-881 . • i,- z laJ L N o 161 t- o 41( t/) laJ ,,, z • I,- _z _<2 ,,,o _"b q, .,-J _") (D _ G. I,- ,3¢ "W I_- I.-. I'-I'-I-,. I--. l--(n Z ° .Ky ° • , . • . . . . . ¢/] 1,,,. • 0 ..,+ +e .+ Z 01 .mr 132 JPL TECHNICAL REPORT NO. 32-881 8 i[ 8 I,- IIi = ee me 9 133 JPL TECHNICAL REPORT NO. 32-881 t.) re. v Y. ,_ x 0 • -J i,J a. i.J 0 eo ! _ o_ =,,==,,=,_ ,, _;-:_; I_ -I _-_ _ o n_ ,( ¢/) O 8 t_ _n ¢J ,( Q go . _ N_Ng_N • °° • ° oo°°°o° r, et N _z • ! Z 134 JPL TECHNICAL REPORT NO. 32-B81 W_ r, @ e@ It m iP a @ 3 _ _ zw _gggG_ _ e@ a¢ t,J i- •-: ,, e_l-- e} A X2' _J • ° • o • • • I1@ x xx ag Ne@ I_lK ,,_r* O N N N ,.., _, L t_ Z ! 7 ! 135 JPL TECHNICAL REPORT NO. 32-881 d" W8 N • .J k_ 8 el Z I ,1¢ 136 JPL TECHNICAL REPORTNO. 32-881 W_ N • ,=4 .= W 8 _ o_ _ -_z _ Z x x x - 8 8 a. =m x 137 JPL TECHNICAL REPORT NO. 32-881 w o. iiiii + + N_ NO • 411 0 i (3, 138 JPL TECHNICAL REPORT NO. 32-881 W_ ,J w w li. S W I¢ J Z N I|1 + O _ o _ < < • _ _ + °'- _. _¢ u-uu _, _ _ O_ L Z _9 "+W IIQ Z + q, 139 JPL TECHNICAL REPORT NO. 32-881 L "_ _ • z •_ • ° •_ ° ° • ° ° ° • ° • ,." S_ _ _8 _ _, • ',. o • q_ 140 JPL TECHNICAL REPORT NO. 32-881 O _m_ W W W W _. g T _ g g O_tl¢ O O O O O O O I$ o _K L 141 JPL TECHNICAL REPORTNO. 32-881 Wc; ..J w le o. S Y I¢ p. ae) • • • • t_ e¢ 4( o ,e_ -Jr O M z _- z ,- _; o. o. ,y. 6 z .z -z _ • _ _ _ .... _ ?,, oo o _xx _W Z 142 ,I JPL TECHNICAL REPORT NO. 32-881 ;8 ill w Ii+ 8 Ib 8 ii- N OO Z a Z +, 143 JPL TECHNICAL REPORT NO. 32-881 • m • L S Y e,. x ¢1¢ 0 8 ,e¢ z z M 9K I-- Z • I-- .KZ eK O i b. c_ z ,: _ _ -- . _ _ ._} _ ' • • • .J eK 1- 0 _ ._. I_ , .., z • "; "; I_,- " L Z 144 JPL TECHNICAL REPORT NO. 32-881 W_ w r, W el If Ih, i,, 00 O0 ! o _L • _ _ **** • ** . . *** o Z M 0 ¢J ,K I- ,r., '*' I[ x .,L 0 _. 0 i 145 JPL TECHNICAL REPORT NO. 32-881 ° W,5 ,.D ,¢,> to 8 Q. S V m i,.. I 4t) F. z la, _. J.J ii'i d Z Z w _ _ . _ _ _. _ ._ • • • • • L 0 _m_ W m m a. 146 JPL TECHNICAL REPORT NO. 32-881 i 8 0 g; o; oo o.o I N_ NZ NZ NZ NZ NZ NZ NZ L Z X 147 JPL TECHNICAL REPORT NO. 32-881 z O i i° o! iii! • ° ° ° ° ° ..... 8 el W N I,,- I¢ N ! °_¢ ,e¢ e¢ Z 148 JPL TECHNICAL REPORT NO. 32-881 • + .. +. • • I_ o • • • • o + t,.. 1o o_ i-i _ i,+ _ +.I ,F'I I"* +., + L V I.. 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Mariner Mars 1964 mechanical configuration

Abstract

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Mariner Mars 1964 mechanical configuration

Abstract

Extracted data

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