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Saturn I Guidance and Control Systems

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This article describes the guidance and control systems used in the ten Saturn I rocckets which were launched in the development program for the operational Saturn IB and Saturn V launch vehicles.
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O
UEST
THE HISTORY OF SPACEFLIGHT QUARTERLY
Volume 17, Number 4
2010
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5In MMemoriam: RRobert TTruax
By Rick W. Sturdevant
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8Deep SSpace NNavigation: The Apollo VVIII Mission
By Paul Cerruzzi
19 Saturn IIGuidance aand Control SSystems
By Edgar Durbin
38 Showing tthe WWay: NASA, tthe NNRO, aand tthe
Apollo LLunar RReconnaissance PProgram
1963—1967
By Vance O. Mitchell
46 Earthshine , Fading
By David Clow
Oral HHistory
32 An IInterview wwith TTh omas CC. DDuxbury
The SSoviet PPhobos aand Russian Mars 996 Missions
Interview by Susan Niebur
57 My EEnc ounter wwith EEdgar MMitchell
By Rick Mulheirn
Contents
Volume 17 • Number 4 2010
Q U E S T 17:4 2010
1
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 its
middle that serves as a spare tire and a bumper. The solar array “para-
sol” suspended above the unicycle rover is fitted with solar batteries to
collect 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, but
in Thompson’s ad the “space motel” mentioned refers to a space station
that would serve as a “motel” for rockets and as a long-term habitation
module for astronauts. Artist: Don Hinkley, Business Week, 12 July 1958
A von Braunian wheel-shaped space station hovers in orbit around the
Moon, a variation on the space station as stepping stone in the form of
an Earth-orbiting workshop for assembly of lunar and interplanetary
spacecraft. Aviation Week, 13 July 1959.
All images courtesy of Blast Books and appear in the book, Another
Science Fiction: Advertising the Space Race 1957—1962.
BBooookk RReevviieewwss
62 On tthe BBookshelf
64 The BBook oof tthe MMoon
Book by Rick Stroud
Review by Roger D. Launius
65 Another SScience FFiction:
Advertising tthe SSpace RRace 11957—1962
Book by Megan Prelinger
Review by Scott Sacknoff
66 Manhattan PProject tto tthe SSanta FFe
Institute: TThe MMemoirs oof GGeorge AA. CCowan
Book by George A. Cowan
Review by Christopher Stone
67 Space EExploration aand HHumanity:
A HHistorical EEncyclopedia
Edited by Stephen Johnson
Review by Scott Sacknoff
Front CCover IImage CCaption aand CCredit
A fish-eye lens view showing astronauts Alan B. Shepard Jr. and
Edgar D. Mitchell in the Apollo lunar module mission simulator
at the Kennedy Space Center during preflight training for the
Apollo XIV lunar landing mission, 15 July 1970. Credit: NASA
Q U E S T 17:4 2010
19
FEATURE
by Edgar Durbin
Introduction
The vehicle that carried Americans to the Moon from
1969–1972 was proposed by the Wernher von Braun group at
Redstone Arsenal to the Department of Defense (DoD) in
December 1957.1The scope of the proposal was broader than a
mission to the Moon: it was “an integrated missile and space
program” for the United States. Its objectives were scientific and
technological supremacy, space warfare capability, and space
exploration.
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 new
Advanced Research and Projects Agency (ARPA) replied to the
Redstone 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 rocket
engines.” The vehicle authorized by ARPA was intended for a
variety of missions, including space defense, large satellite
launch, a troop carrier for missions on Earth, an intercontinental
ballistic missile (ICBM), scientific research, space probes, and
humans in space missions.2Eventually the vehicle was named
Saturn,3and 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 President
John F. Kennedy on 25 May 1961 in a speech at a special joint
session of Congress, by that time, the Saturn had been in devel-
opment for two and a half years. The Saturn program lasted for
18 years, until 1975.4
There were three Saturns: Saturn I, Saturn IB, and Saturn
V. They are compared in Figure 1.
Saturn I was the research-and-development vehicle, and
Saturn IB and Saturn V were the operational vehicles, used to
launch human missions. Figure 1 compares the three vehicles for
a common task, insertion of a payload into low Earth orbit. In
practice, Saturn V performed this mission only once, for launch
of the orbital workshop during the Skylab program. Most of the
Saturn V launches were used for translunar injection during
Apollo missions.6
Saturn I had two configurations, called Block I and Block
II, 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 eight
70-inch propellant tanks surrounding a single 105-inch tank
feeding eight engines. The second stage was the S-IV, which had
a single liquid oxygen tank and a single liquid hydrogen tank
feeding six engines. The instrument unit was the third and last
stage in the Saturn I Block II stack. Above the instrument unit
was the Apollo spacecraft. The complete vehicle was a Saturn
rocket and an Apollo spacecraft. Ten Saturn I vehicles, SA-1
(Saturn–Apollo 1) to SA-10, were launched from Cape
Canaveral Launch Complex 34 (LC-34) and Launch Complex
37B (LC-37B).
Scope
This article describes the guidance and control system
used in the Saturn I vehicles. The Saturn launch vehicle was
steered by swiveling its engines. Actuators could move four of
the eight H-1 engines in the S-I stage through ±7 degrees.8A
similar system of actuators moved the six RL-10 engines of the
S-IV stage through ±4 degrees.9The actuators received com-
mands from the flight control computer, an analog device that
converted data from the digital guidance computer and from
control sensors into actuator commands. The guidance comput-
Saturn I Guidance and Control Systems
Figure 1. Saturn I, IB, and V compared.5Block I Saturn I carried
dummy stages in place of S-IV. Saturn V for Skylab carried an
orbital workshop in place of S-IVB.
Q U E S T 17:4 2010
20
er calculated necessary changes in vehicle
attitude from data received from inertial
sensors carried on the stabilized platform.
The trajectory was determined for early
vehicles by a mechanical cam device and
for later vehicles by a program loaded on
the guidance computer.
The basic guidance system thus
consisted of a program defining the target
trajectory (specified either by the shape of
the cam or by a program stored in the
guidance computer), sensors to measure
vehicle orientation and velocity, and a
digital computer to solve the guidance
equation and calculate changes in attitude
needed to follow the target trajectory.
The control system included angle-
of-attack sensors, attitude rate sensors,
control accelerometers, the analog flight
computer, and actuators to tilt the engines.
System Development
Figure 3 through Figure 7 show the
evolution of the guidance and control sys-
tem from SA-1 through its final version in
SA-10. The major themes of development
were as follows:
· Passenger vs. active—Several compo-
nents first flew on Saturn I as passengers,
not influencing vehicle motions, but only
collecting data that were telemetered to
ground for analysis. In later missions,
some of these components became active;
that is, they provided data to the control
and guidance system.
· Tilt program—All Saturns flew a tra-
jectory that was a function only of time
during passage through the atmosphere,
when aerodynamic forces were greatest.
The pitch of early vehicles was controlled
by a cam device, whose shape constituted
the tilt program. Later vehicles carried a
tilt program in the digital guidance com-
puter as a set of coefficients of a power
series. During this first phase of flight,
guidance sensors only telemetered data to
the ground without influencing the trajec-
tory of the vehicle.
· Roll program—Block I missions did
not have a roll program, though there was
some uncontrolled roll due to thrust
imbalance. Block II missions, which left
from a different launch complex, had to
roll after liftoff to the proper azimuth.
· Digital computer—The digital guid-
ance computer, flown as a passenger on
SA-5 and -6, replaced the cam device (for
the tilt program) and the program device
(for sequence timing), and made active
guidance possible, starting with the sec-
ond stage of SA-6 and continuing on all
subsequent missions. It also made a com-
bined roll–tilt program possible, starting
with SA-7.
· Stabilized platform—The ST-90 stabi-
lized platform, used in Jupiter missiles
and flown on SA-1 to -6, was replaced on
SA-7 by the ST-124, designed for the
Moon mission. (A stabilized platform is a
gyroscopic device for measuring vehicle
velocity and attitude in a space-fixed
coordinate system.)
· Control sensors—Angle-of-attack data
used for control on the first three missions
was replaced by accelerometer data for
the fourth and subsequent missions.
(Angle of attack is the angle between the
vehicle’s longitudinal axis and the air
flow past the vehicle.)
· Instrument Unit—Block I vehicles car-
ried guidance and control sensors in four
pressurized containers housed in the for-
ward part of the S-I stage. These sensors
were carried by Block II vehicles in a new
stage, the Instrument Unit, which stacked
on top of the S-IV stage. SA-5, -6, and -7
carried the first version of the Instrument
Unit; and SA-8, -9, and -10 carried the
second version.
These developments are shown in
Figure 3, which summarizes information
Figure 2. Diagram of Block I and Block II of Saturn I 7
Q U E S T 17:4 2010
21
from more than a dozen technical reports
written during the Saturn program.
Figure 4 to Figure 7 are original
system diagrams from fight evaluation
reports and other contemporary technical
documents edited for clarity.
Control System
The control system sensed devia-
tions from the attitude specified by the
guidance system (pitch, roll, and yaw)
and commanded changes in the direction
of thrust to counter them. The Saturn I
Block I vehicle was unstable during part
of its flight, because the center of pressure
of aerodynamic forces on the vehicle was
forward of its center of mass.15 A small
increase in the angle of attack (for exam-
ple, due to wind) would cause vehicle
rotation, which further increased the
angle of attack. If not corrected, the vehi-
cle would tumble. The addition of fins at
the bottom of the first stage in Block II
improved stability by moving the center
of pressure aft of the center of mass dur-
ing the first minute of flight.16 However,
active control was still necessary and was
employed throughout a mission.
The flight control computer
received outputs from several sensors,
which varied with mission. For the first
three missions, angle of attack measured
by sensors in the nose cone was combined
with data from sensors carried in the ST-
90 stabilized platform. Beginning with
SA-4, control accelerometers were used
instead 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
βis the swivel angle,
φis the error in the vehicle attitude
angle determined by the guidance sys-
tem,
α is the angle of attack, and a0, a1, and
b0are gains, which change during
flight.20
The heading error φwas 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 used
on SA-1 and SA-3.11 Boxes with dashed lines and “Meas. Only!” indicate passenger components.
Q U E S T 17:4 2010
22
by the ST-90 stabilized platform.
Networks of resistors and capacitors in the
control computer calculated the heading
error rate dφ/dt.21
Beginning with SA-4, instead of
angle of attack, the control computer used
lateral accelerations measured by body-
fixed control accelerometers. The control
equation was then
β= a0*φ+ a1*dφ/dt +g2*d2γ/dt2
where d2γ/dt2is the acceleration of the
vehicle normal to its long axis and g2is a
gain.22
Control feedback stability analy-
ses23 and a mathematical condition called
the Drift Minimum Principle (DMP)24
were used to calculate the gains. The
DMP caused the vehicle to head into the
wind slightly, reducing the angle of attack
by about 50 percent, so that the lateral
component of the gimbaled engine thrust
just opposed the lateral aerodynamic
force. In addition to gain (change in
amplitude of the sensor signals), the con-
trol computer introduced a shift in phase
that was critical to achieving stability. The
appropriate phase had been investigated
theoretically by modeling the electrical
circuits in the control system, the mechan-
ical components (actuators, engine gim-
bals), and the vibrations of the vehicle
structure. This system was a closed loop,
and could either tend to the orientation
specified by the guidance computer or
diverge if gain and phase were wrong. In
the latter case, the control system would
over-correct for winds or other random
disturbances and provoke oscillations that
could lead to vehicle breakup. To keep the
system in the region of stability, the sig-
nals from the sensors were passed through
phase shaping networks in the control
computer.
Gains varied during a mission and
from mission to mission. Potentiometers
and resistors switched by relays within the
control computer determined the gains
(a0, a1, and either b0 or g2). A synchro-
nous motor rotated a cam that rotated the
potentiometers. The shape of the cam
grooves and the times when the motor was
turned on and off determined the way that
the gains varied.25 Figure 10 shows the
two methods of setting control gain.
Figure 11 shows the gains used on
SA-2.
SA-2 was controlled only in
response to signals from the ST-90 stabi-
lized platform for the first 25 seconds of
flight. After that time, the gain for the
angle-of-attack signal increased to a max-
imum between 50 and 70 seconds, then
declined to zero again at 90 seconds.
Comparing Figure 11 and Figure 12
shows that the angle-of-attack sensor was
Figure 5. [top] SA-6 Guidance and Control
System12 The system on the left con-
trolled the first stage, and the system on
the right controlled the rest of the mission.
Figure 6. [bottom] SA-7 Guidance and
Control System13 Same as the second
stage guidance used on SA-6: no program
device, ST-90, or tilt cam.
Q U E S T 17:4 2010
23
used during the period of high dynamic
pressure. The sensor was considered
unreliable after about 105 seconds, due to
low dynamic pressure.30 Other missions
showed similar gain profiles. See Figure
13 for the behavior of control gains with
time on the operational Saturn IB vehicle.
Block I Guidance
The guidance systems for Block I
vehicles (SA-1 to -4) were simplified,
since these were suborbital missions, their
azimuths were fixed by the launch pad
configuration, and they had a single pow-
ered stage (S-I). As with all Saturn
launches, the early part of flight, when
aerodynamic forces were greatest, fol-
lowed a preset path, designed to minimize
lateral forces.32 This tilt program was
determined by the shape of a groove in a
cam device housed in the servo loop
amplifier box.33 This cam device is
shown in Figure 14.
Saturn had defined roll, yaw, and
pitch axes (X, Y, and Z, respectively) as
shown in Figure 15. Its preferred roll atti-
tude had Positions I and III in a vertical
plane, with Position I pointing down
toward Earth when the vehicle was hori-
zontal in orbit. The Z axis passed through
Position I (also known as Fin I in Block
II), 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 of
north.35 The Y axis was perpendicular to
X and Z and was horizontal at launch,
pointing roughly south.
The engines were canted to approx-
imately direct their thrust through the
vehicle’s center of gravity. This dimin-
ishes the tendency to rotate caused by an
engine outage or thrust imbalance.
Because of the orientation of LC
34, Block I vehicles could be launched
with no roll maneuver. Figure 16 shows
that this meant that S-band, AZUSA, and
C-band antennas faced downward,
toward the ground tracking stations with
which they communicated. Positions I, II,
III, and IV in Figure 16 are the same as in
Figure 15.
Block II Guidance
The orientation of LC-37 made it
necessary for Block II vehicles to roll
before pitching over in the plane of the
orbit. The amount of roll depended on the
launch azimuth, which was a function of
launch time. The vehicle’s trajectory was
in a plane that contained the center of
Earth, the launch site, and the position of
the Moon at the time of insertion into
Earth 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 (passenger
only)18 Comparing the dynamic pressures at six differ-
ent points on the nose cone allowed calculation of the
direction from which air flowed over the vehicle.
Q U E S T 17:4 2010
24
Figure 9. Local Angle-of-Attack Sensors. These controlled the
flights 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 mechanism
driven by a synchronous motor. The cam positions a potentiometer to
set 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 cost
precious fuel. During countdown, this plane moved, because the
Moon moved, so the roll maneuver and launch azimuth were
functions of the launch time.
The external azimuth alignment system adjusted the orien-
tation of the stabilized platform continuously during countdown
to keep the ξaxis of the stabilized platform pointed at the launch
azimuth. The program device, a precision multitrack tape player
holding sequence data, started at liftoff, and after a delay to allow
the vehicle to clear the launch tower, it signaled the flight
sequencer to close the servo loop, and the roll maneuver began.40
It stopped when the signal from the stabilized platform ended as
the vehicle reached the flight azimuth. Shortly after, the program
device signaled the start of the tilt cam device.
For SA-5 and for the first stage of SA-6, mechanical
devices as described above controlled roll and tilt. From SA-7, a
Q U E S T 17:4 2010
25
digital guidance computer replaced the
program device and the cam tilt
device.41 Roll and tilt maneuvers were
then performed together, rather than
sequentially. Roll was completed within
the first 30 seconds of flight, but pitch
increased continuously almost until
insertion into orbit. Since tilt was calcu-
lated by evaluating a polynomial, the tilt
program was easily changed between
missions by loading the computer with
different coefficients. Reprogramming
the cam tilt device required a month,
because a new cam had to be
machined.42
The digital flight computer made
active guidance possible: guidance
depended on flight conditions, not just
time. The guidance equations were
solved once a second, and the optimum
vehicle attitude was sent to the flight
control computer to create commands
for the engine actuators. Roll, yaw, and
pitch required separate guidance solu-
tions.
Once the correct roll attitude had
been achieved, the flight control system
maintained it fixed. No guidance calcu-
lations were necessary for roll, because
the correct value was determined at
launch and remained constant for the rest
of 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 motor
begins to rotate the ST-90 pitch ring at a rate determined by the shape of
the cam groove. Gyros on the ST-90 stabilized platform sense the rotation
and send signals that cause the vehicle to pitch over in a sense that keeps
the stabilized platform stationary and reduces the gyro signals to zero.
The engines are canted to approx-
imately direct their thrust through
the vehicle’s center of gravity.37
This diminishes the tendency to
rotate caused by an engine outage
or thrust imbalance.
Q U E S T 17:4 2010
26
ed out of the orbit plane. Yaw varied little
from zero, and used the simplest guidance
mode, called delta minimum guidance.
The steering equation for yaw was a lin-
ear function of displacement from the
orbit plane and the time derivative of this
displacement.
The vehicle pitched when it rotated
in the orbit plane from vertical toward the
horizon. Pitch varied more than yaw or
roll, and determining optimum pitch was
a more difficult problem. The first scheme
for pitch guidance flown on Saturn I was
the polynomial guidance mode or PGM,
used on the second stage of SA-6
and the entire powered flight of SA-
7.43 An improved scheme, the iter-
ative guidance mode or IGM, was
used on the rest of Saturn I, IB, and
V missions. PGM and IGM are both
examples of adaptive guidance.
The initial part of a Saturn mis-
sion flew open loop,44 without
active guidance, to reduce aerody-
namic forces during passage
through most of the atmosphere.
The guidance system determined
the desired attitude using only the
time from launch. It did not take account
of variations in thrust, wind, or other dis-
turbances. Active guidance began after
the second stage (S-IV) had ignited, about
165–168 seconds after liftoff.45 At this
point the vehicle altitude was about 90
kilometers, well above the point of maxi-
mum dynamic pressure at 11–12 kilome-
ters. While the position at the ignition of
the first stage was known, because of
variations in flight conditions during the
first-stage burn, the point where active
guidance began might be anywhere in a
considerable volume of space. Afamily of
minimum fuel trajectories was therefore
calculated in the months before launch
which started at different points in this
volume, with various values of velocity
and attitude. Table 2 shows the variations
used to generate the trajectories for SA-7.
An optimum trajectory, with given
starting and ending points, could be deter-
mined using the calculus of variations
with a large digital computer and substan-
tial time. Because it was beyond the capa-
bilities of the onboard digital computer to
use calculus of variations techniques, the
results of computations performed on the
ground were summarized by a polynomi-
al approximation to the optimum steering
angle. The coefficients of the polynomial
were derived by regression analysis of the
set 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 had
about 35 terms. Such polynomials, one
for 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
Q U E S T 17:4 2010
27
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 simplifying
assumptions of a flat Earth and constant gravity field. This allowed
an analytical solution to the guidance equations, which could be
evaluated onboard. By iterative use of this solution (once a sec-
ond), a highly accurate approximation to an optimum (minimum
fuel 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 afforded
another 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.
Conclusion
Saturn I was followed by the Saturn IB and Saturn V pro-
grams, which used the knowledge gained by the Saturn I program
and some of the components. A new instrument unit, the third ver-
sion of that stage, was built by the IBM Federal Systems Division
in 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 stage
for 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 diameter
and 58 inches high, and was both designed and built by NASA
Marshall 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
Q U E S T 17:4 2010
28
er in diameter than version 2 (260 inches
versus 154 inches) but only slightly taller
(36 inches versus 34 inches). While
Version 3 carried the ST-124 stabilized
platform that flew on Saturn I, the ASC-
15 guidance computer was replaced by
the new Launch Vehicle Digital
Computer (LVDC), also made by IBM.
The flight control computer, manufac-
tured by Electronic Communications,
Inc., of St. Petersburg, Florida, had been
a box50 on Saturn I, but for Saturn IB
and V it was a large cylinder.
About the Author
Edgar Durbin is a physicist who, since
retirement from government service in
2002, has worked part time at the
Smithsonian Institution National Air and
Space Museum, Department of Space
History. Most of his research there has
been on navigation, control, and guid-
ance systems used in the Saturn program.
He received a bachelor of arts degree in
mathematics from Harvard University in
1962, a bachelor of arts degree in physics
from Oxford University in 1964, a doc-
torate in physics from Rice University in
1972, and master’s degree in public
administration from the Kennedy School
of Government at Harvard in 1977.
Notes
1. “Proposal: A National Integrated Missile
and Space Vehicle Development Program,”
Army Ballistic Missile Agency, Development
Operations Division, Huntsville, Alabama,
10 December 1957. Report No. D-R-37.
2. Roy Johnson, “ARPA Order No. 14-59,”
Advanced Research Projects Agency, 15
August 1958.
3. Helen T. Wells, Susan H. Whitely, and
Carrie E. Karegeannes, Origins of NASA
Names (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 by
SA-8, -9, and -10, was 154 inches in diameter and 34
inches high. Version 3 flew on Saturn IB and V and
was 260 inches in diameter and 36 inches tall.
Figure 22. ST-124 in the collection of the National Air and
Space Museum.57 This ST-124 was last modified in April 1974,
and may have been manufactured for a mission (Apollo 18 or
19) 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 gimbal
by an external theodolite before launch.
Figure 23. Inertial Coordinate System within the ST-124
Stabilized Platform.58 Three gyros on the inner gimbal sense
any rotations and send signals to circuits outside the ST-124
that generate commands for the servotorque motors that
exactly counter the rotations, stabilizing the inner gimbal in a
fixed orientation. The integrating accelerometers thus meas-
ure vehicle velocity in a fixed (inertial) coordinate system.
Q U E S T 17:4 2010
29
4. The last Saturn vehicle to be launched
was AS-210, on 15 July 1975.
5. David S. Akens, Saturn Illustrated
Chronology (NASA MSFC, 20 January 1971),
MHR-5, appendix F, figure 5, “Saturn Engine
Applications” with additional data added by
the author.
6. For translunar injection, the S-IVB third
stage was reignited to increase spacecraft
velocity to change its orbit around Earth
from approximately circular to highly ellipti-
cal. The timing and length of the burn was
such as to put the spacecraft at apogee
near the Moon. Wikipedia, “Trans Lunar
Injection.” W. David Woods. How Apollo Flew
to the Moon (Springer, 2007), chapter 4,
“Earth Orbit and TLI.”
7. Akens, “Configurations of Saturn Flight
Vehicles,” Saturn Illustrated Chronology, fig-
ure 49.
8. Saturn Flight Evaluation Working Group,
“Saturn SA-1 Flight Evaluation,” NASA
MSFC, 14 December 1961, MPR-SAT-WF-
61-8, 217.
9. B. E. Duran, “Saturn I/IB Launch Vehicle
Operational Status and Experience,” paper
given at Aeronautic and Space Engineering
and Manufacturing Meeting of the Society of
Automotive Engineers, Los Angeles,
California, 7–11 October, 1968, 3.
10. F. W. Brandner, “Technical Information
Summary Concerning Saturn Vehicle SA 2,”
NASA MSFC, memo dated 5 April 1962, TMX
51831, figure 5 with some lines and text
cleaned up for legibility.
11. Saturn Flight Evaluation Working Group,
“Results of the Third Saturn I Launch Vehicle
Test Flight SA 3,” NASA MSFC, 26 February
1964, MSFC MPR-SAT-65-13, 21.
12. Saturn Flight Evaluation Working Group,
“Results of the Sixth Saturn I Launch Vehicle
Test Flight (SA 6),” NASA MSFC, 1 October
1964, MPR-SAT-FE-64-18, figure 7-1.
13. Saturn Flight Evaluation Working Group,
“Results of the Seventh Saturn I Launch
Vehicle Test Flight (SA 7),” NASA MSFC, 30
December 1964, MPR-SAT-FE-64-19, figure
7-1.
14. “Saturn I Electrical Power and Systems
Integration SA 8 through SA 10,” NASA
MSFC, 5 February 1965, NASA TM X-53205,
figure 3, 9.
15. Walter Haeussermann, F. Brooks Moore,
and Gilbert G. Gassaway, “Guidance and
Control Systems for Space Carrier Vehicles,”
Astronautical Engineering and Science from
Peenemünde to Planetary Space. Honoring
the Fiftieth Birthday of Wernher von Braun
(McGraw-Hill, 1963), 166.
16. Ernst D. Geissler and Walter
Haeussermann, “Saturn Guidance and
Control,” American Rocket Society,
Astronautics (February 1962):44, 88.
17. “Results of the Third Saturn I,” 26:
“Statham type control accelerometers will
be flown in ‘closed loop’ control on SA-4, in
place of the local angle-of-attack transduc-
ers.”
18. H. J. Weichel, “SA 8 Flight Test Data
Report,” NASA MSFC, 2 August 1965, NASA
TM X-53308, 8. “Saturn SA 2 Flight
Evaluation,” NASA MSFC, 5 June 1962,
MPR-SAT-WF-62-5, 106.
19. Figure in “Saturn SA-2 Flight
Evaluation,” 297. Description on page 106.
20. “The Apollo ‘A’/Saturn C-1 Launch
Vehicle System,” NASA MSFC, Saturn
Systems Office, 17 July 1961, NASA TM X-
69174, MOR-M-SAT-61-5, 227. Brandner,
“Saturn Vehicle SA-2,” figure 6. There are
actually 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 Space
Vehicles,” IEEE New Links to New Worlds,
1963 National Space Electronics
Symposium, figure 4. Walter
Haeussermann, “Description and
Performance of the Saturn Launch Vehicle’s
Navigation, 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 of
the Third Saturn I,” 23. “Results of the Fifth
Saturn I,” 1.
25. Caudle and Colbert, “Flight Control
Computer,” 267.
26. F. Brooks Moore and Melvin Brooks,
Figure 24.(right) Inertial Coordinate Systems within the ST-90 Stabilized
Platform.59 In the ST-90 the accelerometers on the stabilized platform meas-
ured accelerations in the ξ, η, ζsystem, which was rotated by an angle εwith
respect to the X, Y, Z system. For SA-1 ε= 41 degrees.60 In the ST-124 the
accelerometers 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, the
outermost 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 2010
30
“Saturn Ascending Phase Guidance and
Control Techniques,” Technology of Lunar
Exploration (Academic Press, 1962),
18–209. Figure 8 is on page 206.
27. Rudolf Decher, “The Astrionics System of
Saturn Launch Vehicles,” NASA MSFC,
Huntsville, Alabama, 1 February 1966, NASA
TM X-53384, 35.
28. Brandner, “Saturn Vehicle SA-2,” fig 6.
29. “Saturn SA-2 Flight Evaluation,” 32, fig
4–6.
30. “Saturn SA-2 Flight Evaluation,” 85.
31. “Astrionics System Handbook Saturn
Launch Vehicles,” NASA MSFC, 1 November
1968, MSFC No. IV-4-401-1, IBM No. 68-
966-0002, 3.3–5.
32. Page 1.3–1 of “Astrionics System
Handbook”: “Guidance of the vehicle
throughout the first stage burn is achieved
through 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, figure
6–16.
35. “Project Apollo Coordinate System
Standards,” NASA SE 008-001-1, June 1965.
Figure A-8a, “Saturn I and IB Launch Vehicle
Structural Body Axes.”
36. Brandner, “Saturn Vehicle SA-2,” fig 6.
37. Caudle and Colbert, “Flight Control
Computer,” 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 Vehicle
Test Flight (SA 5),” NASA MSFC, 22
September 1964, MPR-SAT-FE-64-17, 48.
41. A. A. Conway and H. K. Bennett, “Saturn I
Electrical Power and Systems Integration SA
5 through SA 7,” NASA MSFC, Astrionics
Division Electrical Systems Integration
Branch, 25 March 1963, MTP-ASTR-E-63-5,
7: “The control of inflight sequential events
after liftoff is derived from the program
device on SA-5 and SA-6. On SA-7, the control
of sequential events will be derived from the
guidance computer. One channel of the pro-
gram device delivers the pulses to the flight
sequencer.”
42. Brandner, “Saturn Vehicle SA-2,” figure 7:
“Tilt program must be selected about one
month before launch because a change in
program requires a new cam.”
43. R. A. Chapman, “Saturn I Block II
Guidance Summary Report,” NASA MSFC, 23
February 1966, NASA TM X-53398. On page
3 he makes the distinction between guid-
ance for SA-7 and succeeding missions. The
PGM is explicitly described by Walter
Haeussermann, “Guidance and Control of
Saturn Launch Vehicles,” AIAA, Second
Annual 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 a
switch, but by loading a new program in the
guidance 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 only
of time. In active guidance, the heading cal-
culation took account of vehicle position,
velocity and acceleration, which were the
results of the preceding heading commands
plus the influences of atmospheric forces
and vehicle dynamics.
45. Active guidance started for SA-6 at
168.23 seconds after launch; for SA-7 at
165.67 seconds, for SA-8 at 166.69 sec-
onds, and for SA-10 at 168 seconds. See
flight evaluation reports for each mission.
46. Adapted from Table 2.2–1 in
Haeussermann, “Guidance and Control of
Saturn Launch Vehicles,” AIAA Paper 65-304.
47. David H. Schmieder and John B. Winch,
“Adaptive Guidance,” Papers Presented at
Session K of the NASA-University Conference
on the Science and Technology of Space
Exploration, Chicago, Illinois, 1–3 November
1962. NASA-SP-17.
48. Walter Haeussermann and Robert Clifton
Duncan, “Launch Vehicle Inertial Navigation
and Guidance,” Status of Guidance and
Control Methods, Instrumentation, and
Techniques as Applied in the Apollo Project
(NASA, 1964), chapter 2. To be presented at
the lecture series on orbit optimization and
advanced guidance instrumentation,
Advisory Group for Aeronautical Research
and Development, North Atlantic Treaty
Organization, Düsseldorf, Germany, 21–22
October 1964.
49. Haeussermann in AIAA paper 65-304
says for PGM “approximately 100 trajectories
are required.” Also, “The PGM scheme
requires approximately 100 to 300 precalcu-
Figure 27. Saturn I program device.65 This
multitrack precision tape player was used on
both Jupiter and Saturn I to determine the
sequence 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 used
on Titan II and later adapted by NASA for use as the guidance computer
on Saturn I Block II. One of the four side panels is removed, revealing 13
of 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 the
external 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 for
extra support. The plated foil on each cover provides radio interference
shielding. This, in turn, is covered with a protective plastic covering. The
four side covers support a total of fifty-two logic sticks, each containing
four welded encapsulated modules [WEM], consisting of resistor, transis-
tors, capacitors, relays, etc. These logic sticks contain the logic circuitry of
the MGC and are encased inside the side covers of the MGC.”64
Q U E S T 17:4 2010
31
lated trajectories, whereas only one or two
precalculated trajectories are required to pro-
vide the presetting for the IGM scheme.”
Haeussermann and Duncan, “Status of
Guidance” says “The number of precalculat-
ed trajectories required to generate the guid-
ance equations varies from as few as one or
two to as many as 300. The minimum propel-
lant polynomials require the largest number;
the schemes employing a standard reference
trajectory or a set of explicit equations need
the smallest number. Polynomials designed
primarily for accuracy and not for the mini-
mum propellant consumption may require
from 25 to 50 trajectories.”
50. Caudle and Colbert, “Flight Control
Computer,” figure 10, gives no dimensions or
weights. Haeussermann, Moore, and
Gassaway, “Guidance and Control Systems,”
175, says “The control computer used in the
early Saturn C-l vehicles [that is Saturn I] had
a volume of 0.5 ft3, a weight of 30 lb, and a
power requirement of 150 watts. As present-
ly visualized, the Saturn C-5 class control
computer will have no major functional differ-
ences from those previously described.”
“MSFC Artifacts Catalog” shows a “computer
used 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, 15
February 1966, http://hdl.handle.net/
2060/1966 0014308. Courtney G. Brooks,
James M. Grimwood,
and Loyd S. Swenson,
Chariots for Apollo: A
History of Manned
Lunar 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/64
12716.jpg.
57. Photograph taken by Edgar Durbin at
NASM 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 of
Determining the Source of Errors in Guidance
Measurements and the Resultant Errors in
Earth-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 (March
1958): 199–200.
62. Caudle and Colbert, “Flight Control
Computer,” 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, Sheppard
Technical Training Center, May 1967,
Student Study Guide OBR1821F/3121f-V-1
through 4, Volume I of II, 62. Excerpt provid-
ed by Titan Missile Museum, Sahuarita,
Arizona.
65. “MSFC Artifacts Catalogue,” NASA MSFC
Management 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 “The
flight 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 actuation
devices. The basic Saturn flight sequencer provides 10 steps, or distribu-
tion points, for control functions. The basic unit can be connected to one
or 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 to
provide 20 steps or distribution points. Physically, the flight sequencer
consists of relays, diodes, and two printed circuit boards.”
Figure 29. [right] Flight sequences for SA-5, controlled by two tracks
of the program device.67
Article
This computer system is not intended for use in the operation of nuclear facilities, aircraft navigation or communications systems, or air traffic control machines, or for any other uses where the failure of the computer system could lead to death, personal injury, or severe environmental damage.-Apple Computer, Inc., Macbook Users Guide, 2006, p. 109. This month's article emphasizes the role played by reliability in establishing the technology of, and market for, the silicon integrated circuit by citing key developments in the Apollo program.
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Beginning with the challenges presented by Sputnik 1 in 1957, and the formation of NASA, the apollo lunar exploration program is reviewed through Apollo Flight 11. The focal points are the spacecraft including the command and service modules, and the lunar module.
A Method of Determining the Source of Errors in Guidance Measurements and the Resultant Errors in Earth-Fixed Components
  • R A Chapman
R. A. Chapman, "A Method of Determining the Source of Errors in Guidance Measurements and the Resultant Errors in Earth-Fixed Components," NASA MSFC, MTP-AERO-62-76, 22 October 1962, figure 1, 12. 60. "Saturn SA-1 Flight Evaluation," 106.
The New Look in Gimbal Systems
  • F K Mueller
F. K. Mueller, "The New Look in Gimbal Systems," Missiles and Rockets (March 1958): 199-200.
Student Study Guide OBR1821F/3121f-V-1 through 4, Volume I of II, 62. Excerpt provid
  • Colbert Caudle
Caudle and Colbert, "Flight Control Computer," figure 10. 63. Photo provided by Stacy L. Fortner, associate archivist, IBM Corporate Archives, Somers, New York. 64. "Missile Launch/Missile Officer (LGM-25) Missile Systems," USAF, Sheppard Technical Training Center, May 1967, Student Study Guide OBR1821F/3121f-V-1 through 4, Volume I of II, 62. Excerpt provided by Titan Missile Museum, Sahuarita, Arizona. 65. "MSFC Artifacts Catalogue," NASA MSFC Management Services Office, 1 July 1976, 206.