OUESTTHE HISTORY OF SPACEFLIGHT QUARTERLY

Volume 17, Number 42010

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2 Letter ffrom tthe EEditor5 In MMemoriam: RRobert TTruax

By Rick W. Sturdevant

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8 Deep SSpace NNavigation: The Apollo VVIII MissionBy Paul Cerruzzi

19 Saturn II Guidance aand Control SSystemsBy Edgar Durbin

38 Showing tthe WWay: NASA, tthe NNRO, aand tthe Apollo LLunar RReconnaissance PProgram 1963—1967By Vance O. Mitchell

46 Earthshine , FadingBy David Clow

Oral HHistory

32 An IInterview wwith TThomas CC. DDuxburyThe SSoviet PPhobos aand Russian Mars 996 MissionsInterview by Susan Niebur

57 My EEncounter wwith EEdgar MMitchellBy Rick Mulheirn

ContentsVolume 17 • Number 4 2010

Q U E S T 17:4 20101

Back CCover IImage CCaption aand CCredits (clockwise, top left)

Propped by a kickstand while its passengers embark on pedestrian explo-rations, this imaginary rover is a rubber inflatable with a belt around itsmiddle that serves as a spare tire and a bumper. The solar array “para-sol” suspended above the unicycle rover is fitted with solar batteries tocollect and store energy for all the rover’s systems. Artist: Frank Tinsley,Aviation Week, 16 February 1959.

The Lummus Company ad depicts a test chamber on Earth for simulat-ing the lunar environment. Missiles and Rockets, 3 September 1962.

The idea of space tourism would naturally imply a need for lodging, butin Thompson’s ad the “space motel” mentioned refers to a space stationthat would serve as a “motel” for rockets and as a long-term habitationmodule for astronauts. Artist: Don Hinkley, Business Week, 12 July 1958

A von Braunian wheel-shaped space station hovers in orbit around theMoon, a variation on the space station as stepping stone in the form ofan Earth-orbiting workshop for assembly of lunar and interplanetaryspacecraft. Aviation Week, 13 July 1959.

All images courtesy of Blast Books and appear in the book, AnotherScience Fiction: Advertising the Space Race 1957—1962.

BBooookk RReevviieewwss62 On tthe BBookshelf

64 The BBook oof tthe MMoonBook by Rick StroudReview by Roger D. Launius

65 Another SScience FFiction: Advertising tthe SSpace RRace 11957—1962Book by Megan PrelingerReview by Scott Sacknoff

66 Manhattan PProject tto tthe SSanta FFe Institute: TThe MMemoirs oof GGeorge AA. CCowanBook by George A. CowanReview by Christopher Stone

67 Space EExploration aand HHumanity:A HHistorical EEncyclopediaEdited by Stephen JohnsonReview by Scott Sacknoff

Front CCover IImage CCaption aand CCredit

A fish-eye lens view showing astronauts Alan B. Shepard Jr. andEdgar D. Mitchell in the Apollo lunar module mission simulatorat the Kennedy Space Center during preflight training for theApollo XIV lunar landing mission, 15 July 1970. Credit: NASA

Q U E S T 17:4 201019

F E AT U R E

by Edgar Durbin

Introduction

The vehicle that carried Americans to the Moon from1969–1972 was proposed by the Wernher von Braun group atRedstone Arsenal to the Department of Defense (DoD) inDecember 1957.1 The scope of the proposal was broader than amission to the Moon: it was “an integrated missile and spaceprogram” for the United States. Its objectives were scientific andtechnological supremacy, space warfare capability, and spaceexploration.

The proposal came on the heels of the Soviet Union’s suc-cessful orbit of Sputnik I on 4 October 1957, and America’s fail-ure to launch Vanguard on 6 December. The DoD’s newAdvanced Research and Projects Agency (ARPA) replied to theRedstone proposal in October 1958 with funding for develop-ment of “a large space vehicle booster of approximately 1.5 mil-lion pounds thrust based on a cluster of available rocketengines.” The vehicle authorized by ARPA was intended for avariety of missions, including space defense, large satellitelaunch, a troop carrier for missions on Earth, an intercontinentalballistic missile (ICBM), scientific research, space probes, andhumans in space missions.2 Eventually the vehicle was namedSaturn,3 and the missions were limited to humans in space mis-sions (Apollo, Apollo-Soyuz Test Program, and Skylab).Although the Moon landing program was proposed by PresidentJohn F. Kennedy on 25 May 1961 in a speech at a special jointsession of Congress, by that time, the Saturn had been in devel-opment for two and a half years. The Saturn program lasted for18 years, until 1975.4

There were three Saturns: Saturn I, Saturn IB, and SaturnV. They are compared in Figure 1.

Saturn I was the research-and-development vehicle, andSaturn IB and Saturn V were the operational vehicles, used tolaunch human missions. Figure 1 compares the three vehicles fora common task, insertion of a payload into low Earth orbit. Inpractice, Saturn V performed this mission only once, for launchof the orbital workshop during the Skylab program. Most of theSaturn V launches were used for translunar injection duringApollo missions.6

Saturn I had two configurations, called Block I and BlockII, which are compared in Table 1 and Figure 2. A vehicle con-sisted of several stages, stacked one on the other. For example,the first stage of Saturn I Block II was the S-I, a cluster of eight70-inch propellant tanks surrounding a single 105-inch tank

feeding eight engines. The second stage was the S-IV, which hada single liquid oxygen tank and a single liquid hydrogen tankfeeding six engines. The instrument unit was the third and laststage in the Saturn I Block II stack. Above the instrument unitwas the Apollo spacecraft. The complete vehicle was a Saturnrocket and an Apollo spacecraft. Ten Saturn I vehicles, SA-1(Saturn–Apollo 1) to SA-10, were launched from CapeCanaveral Launch Complex 34 (LC-34) and Launch Complex37B (LC-37B).

ScopeThis article describes the guidance and control system

used in the Saturn I vehicles. The Saturn launch vehicle wassteered by swiveling its engines. Actuators could move four ofthe eight H-1 engines in the S-I stage through ±7 degrees.8 Asimilar system of actuators moved the six RL-10 engines of theS-IV stage through ±4 degrees.9 The actuators received com-mands from the flight control computer, an analog device thatconverted data from the digital guidance computer and fromcontrol sensors into actuator commands. The guidance comput-

Saturn I Guidance and Control Systems

Figure 1. Saturn I, IB, and V compared.5 Block I Saturn I carrieddummy stages in place of S-IV. Saturn V for Skylab carried anorbital workshop in place of S-IVB.

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er calculated necessary changes in vehicleattitude from data received from inertialsensors carried on the stabilized platform.The trajectory was determined for earlyvehicles by a mechanical cam device andfor later vehicles by a program loaded onthe guidance computer.

The basic guidance system thusconsisted of a program defining the targettrajectory (specified either by the shape ofthe cam or by a program stored in theguidance computer), sensors to measurevehicle orientation and velocity, and adigital computer to solve the guidanceequation and calculate changes in attitudeneeded to follow the target trajectory.

The control system included angle-of-attack sensors, attitude rate sensors,control accelerometers, the analog flightcomputer, and actuators to tilt the engines.

System DevelopmentFigure 3 through Figure 7 show the

evolution of the guidance and control sys-tem from SA-1 through its final version inSA-10. The major themes of developmentwere as follows:

· Passenger vs. active—Several compo-nents first flew on Saturn I as passengers,not influencing vehicle motions, but onlycollecting data that were telemetered toground for analysis. In later missions,some of these components became active;that is, they provided data to the controland guidance system.

· Tilt program—All Saturns flew a tra-jectory that was a function only of timeduring passage through the atmosphere,when aerodynamic forces were greatest.The pitch of early vehicles was controlled

by a cam device, whose shape constitutedthe tilt program. Later vehicles carried atilt program in the digital guidance com-puter as a set of coefficients of a powerseries. During this first phase of flight,guidance sensors only telemetered data tothe ground without influencing the trajec-tory of the vehicle.

· Roll program—Block I missions didnot have a roll program, though there wassome uncontrolled roll due to thrustimbalance. Block II missions, which leftfrom a different launch complex, had toroll after liftoff to the proper azimuth.

· Digital computer—The digital guid-ance computer, flown as a passenger onSA-5 and -6, replaced the cam device (forthe tilt program) and the program device(for sequence timing), and made activeguidance possible, starting with the sec-ond stage of SA-6 and continuing on allsubsequent missions. It also made a com-bined roll–tilt program possible, startingwith SA-7.

· Stabilized platform—The ST-90 stabi-lized platform, used in Jupiter missilesand flown on SA-1 to -6, was replaced onSA-7 by the ST-124, designed for theMoon mission. (A stabilized platform is agyroscopic device for measuring vehiclevelocity and attitude in a space-fixedcoordinate system.)

· Control sensors—Angle-of-attack dataused for control on the first three missionswas replaced by accelerometer data forthe fourth and subsequent missions.(Angle of attack is the angle between thevehicle’s longitudinal axis and the airflow past the vehicle.)

· Instrument Unit—Block I vehicles car-ried guidance and control sensors in fourpressurized containers housed in the for-ward part of the S-I stage. These sensorswere carried by Block II vehicles in a newstage, the Instrument Unit, which stackedon top of the S-IV stage. SA-5, -6, and -7carried the first version of the InstrumentUnit; and SA-8, -9, and -10 carried thesecond version.

These developments are shown inFigure 3, which summarizes information

Figure 2. Diagram of Block I and Block II of Saturn I 7

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from more than a dozen technical reportswritten during the Saturn program.

Figure 4 to Figure 7 are originalsystem diagrams from fight evaluationreports and other contemporary technicaldocuments edited for clarity.

Control SystemThe control system sensed devia-

tions from the attitude specified by theguidance system (pitch, roll, and yaw)and commanded changes in the directionof thrust to counter them. The Saturn IBlock I vehicle was unstable during partof its flight, because the center of pressureof aerodynamic forces on the vehicle wasforward of its center of mass.15 A smallincrease in the angle of attack (for exam-ple, due to wind) would cause vehiclerotation, which further increased theangle of attack. If not corrected, the vehi-cle would tumble. The addition of fins atthe bottom of the first stage in Block IIimproved stability by moving the centerof pressure aft of the center of mass dur-ing the first minute of flight.16 However,active control was still necessary and wasemployed throughout a mission.

The flight control computerreceived outputs from several sensors,which varied with mission. For the firstthree missions, angle of attack measuredby sensors in the nose cone was combinedwith data from sensors carried in the ST-90 stabilized platform. Beginning withSA-4, control accelerometers were usedinstead of angle-of-attack sensors.17

For SA-1 to -3 the control comput-er calculated the swivel angle of the

engines using the equation

β = a0*φ+ a1*dφ/dt +b0*α where β is the swivel angle, φ is the error in the vehicle attitude

angle determined by the guidance sys-

tem, α is the angle of attack, and a0, a1, andb0 are gains, which change during

flight.20

The heading error φ was supplied

Figure 3. Summary by author of Saturn I guidance and control components

Figure 4. SA-2 Guidance and Control System.10 This is essentially the same as the system usedon SA-1 and SA-3.11 Boxes with dashed lines and “Meas. Only!” indicate passenger components.

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by the ST-90 stabilized platform.Networks of resistors and capacitors in thecontrol computer calculated the headingerror rate dφ/dt.21

Beginning with SA-4, instead ofangle of attack, the control computer usedlateral accelerations measured by body-fixed control accelerometers. The controlequation was then β = a0*φ+ a1*dφ/dt +g2*d2γ/dt2

where d2γ/dt2 is the acceleration of thevehicle normal to its long axis and g2 is a

gain.22Control feedback stability analy-

ses23 and a mathematical condition calledthe Drift Minimum Principle (DMP)24were used to calculate the gains. TheDMP caused the vehicle to head into thewind slightly, reducing the angle of attackby about 50 percent, so that the lateral

component of the gimbaled engine thrustjust opposed the lateral aerodynamicforce.

In addition to gain (change inamplitude of the sensor signals), the con-trol computer introduced a shift in phasethat was critical to achieving stability. Theappropriate phase had been investigatedtheoretically by modeling the electricalcircuits in the control system, the mechan-ical components (actuators, engine gim-bals), and the vibrations of the vehiclestructure. This system was a closed loop,and could either tend to the orientationspecified by the guidance computer ordiverge if gain and phase were wrong. Inthe latter case, the control system wouldover-correct for winds or other randomdisturbances and provoke oscillations thatcould lead to vehicle breakup. To keep thesystem in the region of stability, the sig-nals from the sensors were passed throughphase shaping networks in the controlcomputer.

Gains varied during a mission andfrom mission to mission. Potentiometersand resistors switched by relays within thecontrol computer determined the gains(a0, a1, and either b0 or g2). A synchro-nous motor rotated a cam that rotated thepotentiometers. The shape of the camgrooves and the times when the motor wasturned on and off determined the way thatthe gains varied.25 Figure 10 shows thetwo methods of setting control gain.

Figure 11 shows the gains used onSA-2.

SA-2 was controlled only inresponse to signals from the ST-90 stabi-lized platform for the first 25 seconds offlight. After that time, the gain for theangle-of-attack signal increased to a max-imum between 50 and 70 seconds, thendeclined to zero again at 90 seconds.Comparing Figure 11 and Figure 12shows that the angle-of-attack sensor was

Figure 5. [top] SA-6 Guidance and ControlSystem12 The system on the left con-trolled the first stage, and the system onthe right controlled the rest of the mission.

Figure 6. [bottom] SA-7 Guidance andControl System13 Same as the secondstage guidance used on SA-6: no programdevice, ST-90, or tilt cam.

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used during the period of high dynamicpressure. The sensor was consideredunreliable after about 105 seconds, due tolow dynamic pressure.30 Other missionsshowed similar gain profiles. See Figure13 for the behavior of control gains withtime on the operational Saturn IB vehicle.

Block I GuidanceThe guidance systems for Block I

vehicles (SA-1 to -4) were simplified,since these were suborbital missions, theirazimuths were fixed by the launch padconfiguration, and they had a single pow-ered stage (S-I). As with all Saturnlaunches, the early part of flight, whenaerodynamic forces were greatest, fol-lowed a preset path, designed to minimizelateral forces.32 This tilt program wasdetermined by the shape of a groove in acam device housed in the servo loopamplifier box.33 This cam device isshown in Figure 14.

Saturn had defined roll, yaw, andpitch axes (X, Y, and Z, respectively) asshown in Figure 15. Its preferred roll atti-tude had Positions I and III in a verticalplane, with Position I pointing downtoward Earth when the vehicle was hori-zontal in orbit. The Z axis passed through

Position I (also known as Fin I in BlockII), and at LC-34 it pointed approximate-ly 100 degrees 12 minutes east of north.At LC-37B the Z axis pointed at approxi-mately 90 degrees 12 minutes east ofnorth.35 The Y axis was perpendicular toX and Z and was horizontal at launch,pointing roughly south.

The engines were canted to approx-imately direct their thrust through thevehicle’s center of gravity. This dimin-ishes the tendency to rotate caused by anengine outage or thrust imbalance.

Because of the orientation of LC34, Block I vehicles could be launchedwith no roll maneuver. Figure 16 showsthat this meant that S-band, AZUSA, andC-band antennas faced downward,toward the ground tracking stations withwhich they communicated. Positions I, II,III, and IV in Figure 16 are the same as inFigure 15.

Block II GuidanceThe orientation of LC-37 made it

necessary for Block II vehicles to rollbefore pitching over in the plane of theorbit. The amount of roll depended on thelaunch azimuth, which was a function oflaunch time. The vehicle’s trajectory wasin a plane that contained the center ofEarth, the launch site, and the position ofthe Moon at the time of insertion intoEarth orbit.39 Any other orbital inclina-

Figure 7. SA-8, -9, and -10 Guidance and Control System14

Figure 8. Q-Ball Angle-of-Attack Sensors (passengeronly)18 Comparing the dynamic pressures at six differ-ent points on the nose cone allowed calculation of thedirection from which air flowed over the vehicle.

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Figure 9. Local Angle-of-Attack Sensors. These controlled theflights of SA-1, -2, and -3.19

Figure 11. Control Gains for SA-228

Figure 12. Dynamic Pressure vs Flight Time for SA-2.29

Figure 10. Saturn I Control System26 This figure comes from an arti-cle that describes “Gain Programmer” as “a simple cam mechanismdriven by a synchronous motor. The cam positions a potentiometer toset the gain in each channel.” In the Astrionics System Handbook(1968) this device is called the Control Attenuation Timer (CAT).27

Figure 13. Saturn IB Control Gains.31

tion would require out-of-plane maneuvers, which would costprecious fuel. During countdown, this plane moved, because theMoon moved, so the roll maneuver and launch azimuth werefunctions of the launch time.

The external azimuth alignment system adjusted the orien-tation of the stabilized platform continuously during countdownto keep the ξ axis of the stabilized platform pointed at the launchazimuth. The program device, a precision multitrack tape playerholding sequence data, started at liftoff, and after a delay to allowthe vehicle to clear the launch tower, it signaled the flightsequencer to close the servo loop, and the roll maneuver began.40It stopped when the signal from the stabilized platform ended asthe vehicle reached the flight azimuth. Shortly after, the programdevice signaled the start of the tilt cam device.

For SA-5 and for the first stage of SA-6, mechanicaldevices as described above controlled roll and tilt. From SA-7, a

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digital guidance computer replaced theprogram device and the cam tiltdevice.41 Roll and tilt maneuvers werethen performed together, rather thansequentially. Roll was completed withinthe first 30 seconds of flight, but pitchincreased continuously almost untilinsertion into orbit. Since tilt was calcu-lated by evaluating a polynomial, the tiltprogram was easily changed betweenmissions by loading the computer withdifferent coefficients. Reprogrammingthe cam tilt device required a month,because a new cam had to bemachined.42

The digital flight computer madeactive guidance possible: guidancedepended on flight conditions, not justtime. The guidance equations weresolved once a second, and the optimumvehicle attitude was sent to the flightcontrol computer to create commandsfor the engine actuators. Roll, yaw, andpitch required separate guidance solu-tions.

Once the correct roll attitude hadbeen achieved, the flight control systemmaintained it fixed. No guidance calcu-lations were necessary for roll, becausethe correct value was determined atlaunch and remained constant for the restof the mission.

The vehicle yawed when it point-Figure 15. [top] Body-fixed Coordinates for SA-2.36

Figure 16. [bottom] SA-2 Antennas.38

Figure 14. Tilt Program Cam Device34 At liftoff, the synchronous motorbegins to rotate the ST-90 pitch ring at a rate determined by the shape ofthe cam groove. Gyros on the ST-90 stabilized platform sense the rotationand send signals that cause the vehicle to pitch over in a sense that keepsthe stabilized platform stationary and reduces the gyro signals to zero.

The engines are canted to approx-imately direct their thrust throughthe vehicle’s center of gravity.37This diminishes the tendency torotate caused by an engine outageor thrust imbalance.

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ed out of the orbit plane. Yaw varied littlefrom zero, and used the simplest guidancemode, called delta minimum guidance.The steering equation for yaw was a lin-ear function of displacement from theorbit plane and the time derivative of thisdisplacement.

The vehicle pitched when it rotatedin the orbit plane from vertical toward thehorizon. Pitch varied more than yaw orroll, and determining optimum pitch wasa more difficult problem. The first schemefor pitch guidance flown on Saturn I wasthe polynomial guidance mode or PGM,

used on the second stage of SA-6and the entire powered flight of SA-7.43 An improved scheme, the iter-ative guidance mode or IGM, wasused on the rest of Saturn I, IB, andV missions. PGM and IGM are bothexamples of adaptive guidance.

The initial part of a Saturn mis-sion flew open loop,44 withoutactive guidance, to reduce aerody-namic forces during passagethrough most of the atmosphere.The guidance system determinedthe desired attitude using only the

time from launch. It did not take accountof variations in thrust, wind, or other dis-turbances. Active guidance began afterthe second stage (S-IV) had ignited, about165–168 seconds after liftoff.45 At thispoint the vehicle altitude was about 90kilometers, well above the point of maxi-mum dynamic pressure at 11–12 kilome-ters. While the position at the ignition ofthe first stage was known, because ofvariations in flight conditions during thefirst-stage burn, the point where activeguidance began might be anywhere in aconsiderable volume of space. A family of

minimum fuel trajectories was thereforecalculated in the months before launchwhich started at different points in thisvolume, with various values of velocityand attitude. Table 2 shows the variationsused to generate the trajectories for SA-7.

An optimum trajectory, with givenstarting and ending points, could be deter-mined using the calculus of variationswith a large digital computer and substan-tial time. Because it was beyond the capa-bilities of the onboard digital computer touse calculus of variations techniques, theresults of computations performed on theground were summarized by a polynomi-al approximation to the optimum steeringangle. The coefficients of the polynomialwere derived by regression analysis of theset of optimum trajectories.47 The vari-ables of the polynomial were the quanti-ties measured by onboard devices: posi-tion, velocity, and acceleration. The poly-nomial was of the third power, and hadabout 35 terms. Such polynomials, onefor pitch and one for yaw steering angle,were evaluated once a second, with updat-ed values of position, velocity, and accel-eration. This scheme for guidance was

46

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called PGM, and was used only on SA-6 and SA-7.An improved scheme, IGM, was first tested on SA-9, refined

on SA-8 and SA-10, and used on all subsequent Saturn launches.Derivation of IGM steering equations began with simplifyingassumptions of a flat Earth and constant gravity field. This allowedan analytical solution to the guidance equations, which could beevaluated onboard. By iterative use of this solution (once a sec-ond), a highly accurate approximation to an optimum (minimumfuel consumption) trajectory could be achieved.

The advantages of IGM over Polynomial Guidance48 were:

1. Reduced preflight computation. The number of trajectories con-sidered was reduced by two orders of magnitude.49 This affordedanother advantage:

2. Greater flexibility. Fewer preflight calculations could be per-formed more quickly, allowing easier adjustment to changed mis-sion conditions.

3. Greater accuracy and hence improved fuel economy.

ConclusionSaturn I was followed by the Saturn IB and Saturn V pro-

grams, which used the knowledge gained by the Saturn I programand some of the components. A new instrument unit, the third ver-sion of that stage, was built by the IBM Federal Systems Divisionin Huntsville for Saturn IB and V. Version 3 was considerably larg-

Figure 17. Guidance components were carried in four pressur-ized canisters stowed in the forward part of the S-I first stagefor Saturn I Block I.52

Figure 19. [above]Mockup of version 1 of the Instrument Unit, to be flown on SA-5.A separate stage, the Instrument Unit, carried Block II guid-ance components. The cylinder was 154 inches in diameterand 58 inches high, and was both designed and built by NASAMarshall Space Flight Center (MSFC).54

Figure 20. [left]Exploded view of version 1 of the Instrument Unit, with com-ponents in pressurized tubes.55

Figure 18. Canisters are installed on the S-I. 53

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er in diameter than version 2 (260 inchesversus 154 inches) but only slightly taller(36 inches versus 34 inches). WhileVersion 3 carried the ST-124 stabilizedplatform that flew on Saturn I, the ASC-15 guidance computer was replaced bythe new Launch Vehicle DigitalComputer (LVDC), also made by IBM.The flight control computer, manufac-tured by Electronic Communications,Inc., of St. Petersburg, Florida, had beena box50 on Saturn I, but for Saturn IBand V it was a large cylinder.

About the AuthorEdgar Durbin is a physicist who, sinceretirement from government service in2002, has worked part time at theSmithsonian Institution National Air andSpace Museum, Department of SpaceHistory. Most of his research there hasbeen on navigation, control, and guid-ance systems used in the Saturn program.He received a bachelor of arts degree inmathematics from Harvard University in1962, a bachelor of arts degree in physicsfrom Oxford University in 1964, a doc-torate in physics from Rice University in1972, and master’s degree in public

administration from the Kennedy Schoolof Government at Harvard in 1977.

Notes1. “Proposal: A National Integrated Missileand Space Vehicle Development Program,”Army Ballistic Missile Agency, DevelopmentOperations Division, Huntsville, Alabama,10 December 1957. Report No. D-R-37.

2. Roy Johnson, “ARPA Order No. 14-59,”Advanced Research Projects Agency, 15August 1958.

3. Helen T. Wells, Susan H. Whitely, andCarrie E. Karegeannes, Origins of NASANames (NASA Scientific and Technical lnfor-mation Office, 1976), NASA SP-4402, 17.

Figure 21. Version 2 of the Instrument Unit in the fore-ground and Version 3 behind.56 Version 2, carried bySA-8, -9, and -10, was 154 inches in diameter and 34inches high. Version 3 flew on Saturn IB and V andwas 260 inches in diameter and 36 inches tall.

Figure 22. ST-124 in the collection of the National Air andSpace Museum.57 This ST-124 was last modified in April 1974,and may have been manufactured for a mission (Apollo 18 or19) that was canceled. A temperature-controlled, methanol-water mixture flowing through coils on the top and bottom cov-ers held the temperature of the ST-124 constant. A window,facing down to the right, allowed alignment of the inner gimbalby an external theodolite before launch.

Figure 23. Inertial Coordinate System within the ST-124Stabilized Platform.58 Three gyros on the inner gimbal senseany rotations and send signals to circuits outside the ST-124that generate commands for the servotorque motors thatexactly counter the rotations, stabilizing the inner gimbal in afixed orientation. The integrating accelerometers thus meas-ure vehicle velocity in a fixed (inertial) coordinate system.

Q U E S T 17:4 201029

4. The last Saturn vehicle to be launchedwas AS-210, on 15 July 1975.

5. David S. Akens, Saturn IllustratedChronology (NASA MSFC, 20 January 1971),MHR-5, appendix F, figure 5, “Saturn EngineApplications” with additional data added bythe author.

6. For translunar injection, the S-IVB thirdstage was reignited to increase spacecraftvelocity to change its orbit around Earthfrom approximately circular to highly ellipti-cal. The timing and length of the burn wassuch as to put the spacecraft at apogeenear the Moon. Wikipedia, “Trans LunarInjection.” W. David Woods. How Apollo Flewto the Moon (Springer, 2007), chapter 4,“Earth Orbit and TLI.”

7. Akens, “Configurations of Saturn FlightVehicles,” Saturn Illustrated Chronology, fig-ure 49.

8. Saturn Flight Evaluation Working Group,“Saturn SA-1 Flight Evaluation,” NASAMSFC, 14 December 1961, MPR-SAT-WF-61-8, 217.

9. B. E. Duran, “Saturn I/IB Launch VehicleOperational Status and Experience,” papergiven at Aeronautic and Space Engineeringand Manufacturing Meeting of the Society ofAutomotive Engineers, Los Angeles,California, 7–11 October, 1968, 3.

10. F. W. Brandner, “Technical InformationSummary Concerning Saturn Vehicle SA 2,”NASA MSFC, memo dated 5 April 1962, TMX51831, figure 5 with some lines and text

cleaned up for legibility.

11. Saturn Flight Evaluation Working Group,“Results of the Third Saturn I Launch VehicleTest Flight SA 3,” NASA MSFC, 26 February1964, MSFC MPR-SAT-65-13, 21.

12. Saturn Flight Evaluation Working Group,“Results of the Sixth Saturn I Launch VehicleTest Flight (SA 6),” NASA MSFC, 1 October1964, MPR-SAT-FE-64-18, figure 7-1.

13. Saturn Flight Evaluation Working Group,“Results of the Seventh Saturn I LaunchVehicle Test Flight (SA 7),” NASA MSFC, 30December 1964, MPR-SAT-FE-64-19, figure7-1.

14. “Saturn I Electrical Power and SystemsIntegration SA 8 through SA 10,” NASAMSFC, 5 February 1965, NASA TM X-53205,figure 3, 9.

15. Walter Haeussermann, F. Brooks Moore,and Gilbert G. Gassaway, “Guidance andControl Systems for Space Carrier Vehicles,”Astronautical Engineering and Science fromPeenemünde to Planetary Space. Honoringthe Fiftieth Birthday of Wernher von Braun(McGraw-Hill, 1963), 166.

16. Ernst D. Geissler and WalterHaeussermann, “Saturn Guidance andControl,” American Rocket Society,Astronautics (February 1962):44, 88.

17. “Results of the Third Saturn I,” 26:“Statham type control accelerometers willbe flown in ‘closed loop’ control on SA-4, inplace of the local angle-of-attack transduc-ers.”

18. H. J. Weichel, “SA 8 Flight Test DataReport,” NASA MSFC, 2 August 1965, NASATM X-53308, 8. “Saturn SA 2 FlightEvaluation,” NASA MSFC, 5 June 1962,MPR-SAT-WF-62-5, 106.

19. Figure in “Saturn SA-2 FlightEvaluation,” 297. Description on page 106.

20. “The Apollo ‘A’/Saturn C-1 LaunchVehicle System,” NASA MSFC, SaturnSystems Office, 17 July 1961, NASA TM X-69174, MOR-M-SAT-61-5, 227. Brandner,“Saturn Vehicle SA-2,” figure 6. There areactually three equations, one each for pitch,yaw, and roll.

21. Brandner, “Saturn Vehicle SA-2,” 2.

22. John M. Caudle and Donald C. Colbert,“Flight Control Computer for Saturn SpaceVehicles,” IEEE New Links to New Worlds,1963 National Space ElectronicsSymposium, figure 4. WalterHaeussermann, “Description andPerformance of the Saturn Launch Vehicle’sNavigation, Guidance, and Control System.”NASA MSFC, Huntsville, Alabama, July 1970,NASA TN D-5869, 10.

23. “Apollo A/Saturn C-1,” section 14,“Control Dynamics.”

24. “Apollo A/Saturn C-1,” 227. “Results ofthe Third Saturn I,” 23. “Results of the FifthSaturn I,” 1.

25. Caudle and Colbert, “Flight ControlComputer,” 267.

26. F. Brooks Moore and Melvin Brooks,

Figure 24.(right) Inertial Coordinate Systems within the ST-90 StabilizedPlatform.59 In the ST-90 the accelerometers on the stabilized platform meas-ured accelerations in the ξ, η, ζ system, which was rotated by an angle ε withrespect to the X, Y, Z system. For SA-1 ε = 41 degrees.60 In the ST-124 theaccelerometers measured accelerations in the same X, Y, Z system as the sta-bilized platform. The gimbal system in the ST-124 was “external.” That is, theoutermost gimbal was attached to the vehicle. In the ST-90 the innermost gim-bal was attached to the vehicle, an arrangement called “internal.”61

Figure 25. Saturn I flight control computer.62

Q U E S T 17:4 201030

“Saturn Ascending Phase Guidance andControl Techniques,” Technology of LunarExploration (Academic Press, 1962),18–209. Figure 8 is on page 206.

27. Rudolf Decher, “The Astrionics System ofSaturn Launch Vehicles,” NASA MSFC,Huntsville, Alabama, 1 February 1966, NASATM X-53384, 35.

28. Brandner, “Saturn Vehicle SA-2,” fig 6.

29. “Saturn SA-2 Flight Evaluation,” 32, fig4–6.

30. “Saturn SA-2 Flight Evaluation,” 85.

31. “Astrionics System Handbook SaturnLaunch Vehicles,” NASA MSFC, 1 November1968, MSFC No. IV-4-401-1, IBM No. 68-966-0002, 3.3–5.

32. Page 1.3–1 of “Astrionics SystemHandbook”: “Guidance of the vehiclethroughout the first stage burn is achievedthrough a preset time-tilt (pitch) program.”

33. Brandner, “Saturn Vehicle SA-2,” 2.

34. Brandner, “Saturn Vehicle SA-2,” figure 7.See also Saturn SA-2 Flight Evaluation, figure6–16.

35. “Project Apollo Coordinate System

Standards,” NASA SE 008-001-1, June 1965.Figure A-8a, “Saturn I and IB Launch VehicleStructural Body Axes.”

36. Brandner, “Saturn Vehicle SA-2,” fig 6.

37. Caudle and Colbert, “Flight ControlComputer,” 1.

38. A tracing of Brandner, “Saturn Vehicle SA-2,” figure 3

39. Decher, “Astrionics System of Saturn,”25.

40. Saturn Flight Evaluation Working Group,“Results of the Fifth Saturn I Launch VehicleTest Flight (SA 5),” NASA MSFC, 22September 1964, MPR-SAT-FE-64-17, 48.

41. A. A. Conway and H. K. Bennett, “Saturn IElectrical Power and Systems Integration SA5 through SA 7,” NASA MSFC, AstrionicsDivision Electrical Systems IntegrationBranch, 25 March 1963, MTP-ASTR-E-63-5,7: “The control of inflight sequential eventsafter liftoff is derived from the programdevice on SA-5 and SA-6. On SA-7, the controlof sequential events will be derived from theguidance computer. One channel of the pro-gram device delivers the pulses to the flightsequencer.”

42. Brandner, “Saturn Vehicle SA-2,” figure 7:“Tilt program must be selected about onemonth before launch because a change inprogram requires a new cam.”

43. R. A. Chapman, “Saturn I Block IIGuidance Summary Report,” NASA MSFC, 23February 1966, NASA TM X-53398. On page3 he makes the distinction between guid-ance for SA-7 and succeeding missions. ThePGM is explicitly described by WalterHaeussermann, “Guidance and Control ofSaturn Launch Vehicles,” AIAA, SecondAnnual Meeting, San Francisco, California,26–29 July 1965, paper 65-304.

44. “Open loop” connotes a feedback loop

that is not closed. Changing to active guid-ance was not accomplished by closing aswitch, but by loading a new program in theguidance computer that changed the algo-rithm by which the desired heading was cal-culated. In open loop the heading was calcu-lated by a polynomial that was a function onlyof time. In active guidance, the heading cal-culation took account of vehicle position,velocity and acceleration, which were theresults of the preceding heading commandsplus the influences of atmospheric forcesand vehicle dynamics.

45. Active guidance started for SA-6 at168.23 seconds after launch; for SA-7 at165.67 seconds, for SA-8 at 166.69 sec-onds, and for SA-10 at 168 seconds. Seeflight evaluation reports for each mission.

46. Adapted from Table 2.2–1 inHaeussermann, “Guidance and Control ofSaturn Launch Vehicles,” AIAA Paper 65-304.

47. David H. Schmieder and John B. Winch,“Adaptive Guidance,” Papers Presented atSession K of the NASA-University Conferenceon the Science and Technology of SpaceExploration, Chicago, Illinois, 1–3 November1962. NASA-SP-17.

48. Walter Haeussermann and Robert CliftonDuncan, “Launch Vehicle Inertial Navigationand Guidance,” Status of Guidance andControl Methods, Instrumentation, andTechniques as Applied in the Apollo Project(NASA, 1964), chapter 2. To be presented atthe lecture series on orbit optimization andadvanced guidance instrumentation,Advisory Group for Aeronautical Researchand Development, North Atlantic TreatyOrganization, Düsseldorf, Germany, 21–22October 1964.

49. Haeussermann in AIAA paper 65-304says for PGM “approximately 100 trajectoriesare required.” Also, “The PGM schemerequires approximately 100 to 300 precalcu-

Figure 27. Saturn I program device.65 Thismultitrack precision tape player was used onboth Jupiter and Saturn I to determine thesequence of certain time-dependent events.13 inches by 11 inches by 7.5 inches.

Figure 26. ASC-15 digital guidance computer.63 This computer was usedon Titan II and later adapted by NASA for use as the guidance computeron Saturn I Block II. One of the four side panels is removed, revealing 13of the 52 logic sticks. The port in the middle of the top is for cooling air.“The outer frame consists of four aluminum corner posts which are weld-ed to, and separated by, a series of aluminum rails to form a boxlike struc-ture. Four identical side covers, plus a bottom and top cover, constitute theexternal appearance of the MGC [Missile Guidance Computer, or ASC-15].All covers are formed from laminated plastic, covered with gold-plated alu-minum foil. The covers are convex and are formed with stiffening ridges forextra support. The plated foil on each cover provides radio interferenceshielding. This, in turn, is covered with a protective plastic covering. Thefour side covers support a total of fifty-two logic sticks, each containingfour welded encapsulated modules [WEM], consisting of resistor, transis-tors, capacitors, relays, etc. These logic sticks contain the logic circuitry ofthe MGC and are encased inside the side covers of the MGC.”64

Q U E S T 17:4 201031

lated trajectories, whereas only one or twoprecalculated trajectories are required to pro-vide the presetting for the IGM scheme.”Haeussermann and Duncan, “Status ofGuidance” says “The number of precalculat-ed trajectories required to generate the guid-ance equations varies from as few as one ortwo to as many as 300. The minimum propel-lant polynomials require the largest number;the schemes employing a standard referencetrajectory or a set of explicit equations needthe smallest number. Polynomials designedprimarily for accuracy and not for the mini-mum propellant consumption may requirefrom 25 to 50 trajectories.”

50. Caudle and Colbert, “Flight ControlComputer,” figure 10, gives no dimensions orweights. Haeussermann, Moore, andGassaway, “Guidance and Control Systems,”175, says “The control computer used in theearly Saturn C-l vehicles [that is Saturn I] hada volume of 0.5 ft3, a weight of 30 lb, and apower requirement of 150 watts. As present-ly visualized, the Saturn C-5 class controlcomputer will have no major functional differ-ences from those previously described.”“MSFC Artifacts Catalog” shows a “computerused on the first flour flights of SATURN I”that is 24 inches by 12 1/8 inches x 12 inch-es [that is 2.0 ft3 ] and weighs 55 pounds.

51. “Saturn I Summary.” NASA MSFC, 15February 1966, http://hdl.handle.net/2060/1966 0014308. Courtney G. Brooks,

James M. Grimwood,and Loyd S. Swenson,Chariots for Apollo: AHistory of MannedLunar Spacecraft.NASA History Series,SP-4205, 1979.

52. “Steps to Saturn,”NASA MSFC, 29.http://hdl.handle.net/2060/19660083255.

53. “Steps to Saturn,” 29.

54. “Steps to Saturn,” 66.

55. “ApolloA/Saturn C-1,” figure 6.2, 133.

56. NASA MSFC Image Exchange (MIX).http://mix.msfc.nasa.gov/IMAGES/HIGH/6412716.jpg.

57. Photograph taken by Edgar Durbin atNASM Garber Facility, 4 February 2010.

58. “Saturn V Flight Manual SA-507,” MSFC-MAN-507, 15 August 1969, figure 7-16,7–18.

59. R. A. Chapman, “A Method ofDetermining the Source of Errors in GuidanceMeasurements and the Resultant Errors inEarth-Fixed Components,” NASA MSFC, MTP-AERO-62-76, 22 October 1962, figure 1, 12.

60. “Saturn SA-1 Flight Evaluation,” 106.

61. F. K. Mueller, “The New Look in Gimbal

Systems,” Missiles and Rockets (March1958): 199–200.

62. Caudle and Colbert, “Flight ControlComputer,” figure 10.63. Photo provided by Stacy L. Fortner, asso-ciate archivist, IBM Corporate Archives,Somers, New York.

64. “Missile Launch/Missile Officer (LGM-25) Missile Systems,” USAF, SheppardTechnical Training Center, May 1967,Student Study Guide OBR1821F/3121f-V-1through 4, Volume I of II, 62. Excerpt provid-ed by Titan Missile Museum, Sahuarita,Arizona.

65. “MSFC Artifacts Catalogue,” NASA MSFCManagement Services Office, 1 July 1976,206.

66. Conway and Bennett, “SA-5 through SA-7,” 38–40.

67. Conway and Bennett, “SA-5 through SA-7,” 12.

Figure 28. [Above] Flight sequencer (bottom) and slave (top).66 “Theflight sequencer is a relay device that functions as a step switch to pro-gram distribution of 28 V d.c. power to relays and other control actuationdevices. The basic Saturn flight sequencer provides 10 steps, or distribu-tion points, for control functions. The basic unit can be connected to oneor more slave units to increase the number of steps in multiples of ten.For Saturn use, the flight sequencer is used with a single slave unit toprovide 20 steps or distribution points. Physically, the flight sequencerconsists of relays, diodes, and two printed circuit boards.”

Figure 29. [right] Flight sequences for SA-5, controlled by two tracksof the program device.67