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A Hybrid Simulation for Dynamic Verif icatio
of Saturn Guidance and Control Subsystem:
�IBM NO. 68-U6O-0013
A HYBRID SIMULATION FOR DYNAMIC VERIFICATION O F
SATURN GUIDANCE AND CONTROL SUBSYSTEMS
.
Ronald T Patray
May 15, 1968
International Business Machines Corporation
Federal Systems Division
Space Systems Center
Huntsville, Alabama
�A HYBRID SIMULATION FOR DYNAMIC VERIFICATION O F
SATURN GUIDANCE AND CONTROL SUBSYSTEMS
Ronald T. Patray
International Business Machines Corporation
Federal Systems Division
Space Systems Center
Huntsville , Alabama
I. INTRODUCTION
This paper presents a discussion of a hybrid simulation used to dynamically verify the Saturn Guidance and Control subsystems. First, the Saturn
vehicle is briefly described to provide background information. The Instrument Unit (IU) is considered in more detail to give a proper setting f o r the
Guidance and Flight Control (G and FC) discussion that follows. After a brief
description of the actual G and FC System operation, simulation models of the
G and FC components a r e considered in detail. This is followed by a discussion of the model assignment to a particular computer (digital o r analog) and
justification f o r making that assignment. Finally, results of the A S - 2 0 4 / ~ ~ 1
hybrid simulation studies a r e briefly considered with mention of the actual
flight data. 1
11. SATURN VEHICLE DESCRIPTION
The Saturn IB, which has two propulsive stages (Slide I ) , is serving
as a launch vehicle f o r the Apollo spacecraft earth orbital flight tests. These
flights simulate certain studies of the lunar landing mission and provide flight
1
Some of the material in this paper is based on notes f r o m an IBM conference
presentation entitled, "Digital Computer Program for Support of Hybrid
Computer Simulation of Saturn Launch Vehicle, " by E. W. Snyder.
�t e s t s for the spacecraft and the S-IVB/IU Stage. Each Saturn IB has
a payload consisting of some combination of a Lunar Module (LM), Service
Module (SM), Command Module (CM), and Launch Escape System (LES).
The Saturn V, which has three stages (Slide I ) , is the launch
vehicle f o r the actual lunar landing missions. For these missions the Saturn V
will c a r r y a payload consisting of an LM, SM, CM, and LES.
111. INSTRUMENT UNIT
The major purpose of the Instrument Unit (IU) shown in Slide 2 is to provide the Saturn vehicle with a centralized astrionics package for guidance,
control, sequencing, and telemetry during boost and earth orbit, and through
lunar trajectory insertion. The IU subsystems (Slide 3 ) include the Structural
Portion, Guidance, Flight Control, Environmental, Instrumentation, and
Electrical.
A. Guidance and Flight Control Subsvstems
Slide 4 lists the major components of the Guidance and Flight Control
Subsystems. Included a r e the Launch Vehicle Digital Computer (LVDC), the
Flight Control Computer (FCC), the ST-124 Inertial Platform, and the control
accelerometer and rate gyro /control signal processor.
Slide 5 shows a rough sketch depicting the closed loop operation of the
Guidance and Flight Control Subsystems. Sensors on the inertial platform meas u r e the angles ( 8 ) between the inertial and body reference f r a m e s and changes
in velocity along each inertial axis. These signals a r e transmitted to the
LVDC-LVDA where the velocities a r e used in navigation and in computing the
commanded gimbal angles (x). The actual platform gimbal angles ( 8 ) a r e differenced
with the x's to give attitude e r r o r signals ($1 in the inertial platform frame.
These $ 's a r e transformed to result in attitude e r r o r signals ( A qb ) in the body
frame, which a r e fed to the FCC where they a r e filtered and summed with the
filtered rate gyro signals
to give engine actuator commands (4).The ,f3
signals drive the actuators and thus change the thrust vector orientation which
in turn changes the vehicle attitude.
(6)
IV.
DYNAMIC VERIFICATION OF THE GUIDANCE AND CONTROL SYSTEM
Filters in the FCC and parameters in the LVDC flight program a r e
designed to give satisfactory stability margins while maintaining good vehicle
�response to guidance commands. These designs a r e determined by using
linearized models and linear stability analysis techniques at a few frozen
points in time. While this method of design has proved, thus f a r , to be reliable, it does not consider the effects of nonlinearities in the guidance system,
nor does it consider vehicle dynamics continuously throughout boost flight.
Thus, a means of dynamically checking the FCC filter design and LVDC flight
program in a total system configuration under flight conditions, over all times
of vehicle boost flight, was needed to ensure the nonexistence of adverse
dynamic effects on the vehicle, the astronauts, and the guidance accuracy. It
was decided that a system would be devised in which, except for the LVDC
flight program, all components significantly affecting the vehicle dynamics
would be simulated. For this task, a six degree-of -f reedom real time hybrid
simulation with LVDC/LVDA flight hardware in the loop was chosen that would
fully exercise the flight program and check its dynamic effects on the vehicle
throughout simulated boost flight, and at the same time check the FCC filter
design for vehicle stability and response.
V. HYBRID SIMULATION
Slide 6 shows a simplified G & FC loop for pitch and yaw. The LVDC and
LVDA a r e flight-type hardware, while the Flight Control Computer, Engine
Actuator, Vehicle Dynamics, Rate Gyro, Inertial Platform Assembly, and
Propellant Utilization System (PU) a r e simulated on the hybrid system. Each
component that significantly contributes to vehicle dynamics is described
below, followed by a description of the simulation model of each component.
A. Launch Vehicle Digital Computer (LVDC)
1. Actual LVDC
Slide 7 depicts the Saturn V flight computer (LVDC) tasks by phase:
Phase I includes the time of f i r s t stage boost. During this phase the
vehicle is moving through the dense portion of the atmosphere where high
aerodynamic pressure occurs. To avoid excessive structural loads caused
by guidance maneuvers, no guidance constraints a r e applied. An open loop
guidance scheme in the form of a time tilt program is used.
Phase I1 includes the second stage boost time. The flight program
during this phase uses a path adaptive guidance scheme called Iterative
�Guidance Mode (IGM) in which guidance is a function of space-fixed position (F )
and velocity (G ) , ~ / m
and time. This adaptive guidance scheme seeks to
attain predetermined space -fixed position and velocity vectors with the consumption of a minimum amount of propellant.
Phase I11 includes the f i r s t burn time of third stage boost and orbital
time. IGM guidance is used during the boost portion of this phase.
Phase IV includes the second burn time of third stage boost. IGM
guidance is also used during this phase.
Phase V includes all mission time after $-NB/IU Apollo Spacecraft
separation.
The LVDC/LVDA has a s inputs inertial platform gimbal angles, measured changes in velocity along each platform axis, and flight sequencing
discretes. The LVDC/LVDA outputs a r e attitude-error steering commands
and flight sequencing discretes.
Since an actual flight type LVDC/LVDA is used in the simulation, no
LVDC/LVDA model was devised.
B. ST- 124 Inertial Platform Assemblv
1. Actual Platform
The Inertial Platform Assembly (Slide 6) is the main sensor for guidance.
At guidance reference release, which occurs a few seconds before liftoff, the
platform becomes inertial (space-fixed). A resolver attached to each platform
surface measures the gimbal angle ( 8 ) between the surface and vehicle axis.
These resolver outputs a r e read by the LVDC flight program every 40 milliseconds.
Integrating accelerometers, mounted along each axis of the platform,
sense changes in measured velocities (AX,, A * ~ , A 2,) in increments of
.05 m/sec. Signals representing these velocity changes automatically increment o r decrement velocity counters in the LVDA. These counters a r e read
periodically by the LVDC flight program.
�2.
Platform Simulation Model
Platform gimbal angles ( 0 ) a r e simulated by first transforming the
simulated body r a t e s
(Slide 6) into the inertial coordinate frame, and then
integrating the resulting gimbal angle r a t e s (8). Gimbal angles a r e computed
and transmitted to the LVDA at a 40-millisecond rate.
(4)
The platform integrating accelerometers a r e simulated in two ways:
During f i r s t stage simulation, vehicle thrust is obtained through a table lookup
scheme of thrust versus vehicle altitude. Thrust is then used with remaining
vehicle m a s s (m,), which is computed, to obtain total vehicle acceleration
.
engine angles
(F/mV). This acceleration is resolved through the simulated
( P ) to result in body acceleration components gB,yB, ZB). which in turn
a r e transformed via the platform gimbal angles (0) into inertial platform
acceleration components Wm, Y,,
z,). These accelerations a r e then used
to compute changes in measured platform velocity components
(AX,,
A
A i m ) which would normally be sensed by the platform integrating accelerometers. The measured velocity changes a r e computed and
transmitted through special interface equipment to the LVDA every 40 milliseconds. During the second and third stage simulation, vehicle thrust is
obtained through a table lookup scheme of thrust versus Propellant Utilization
System (PU) valve position, which is the output of the PU System simulation
model. After vehicle thrust has been obtained, changes in the platformmeasured velocities (AX,, A
, A z,) a r e derived in the same manner as
in the f i r s t stage simulation.
.
C.
Flight Control Computer (FCC)
1. Actual FCC
The primary functions of the FCC a r e to provide command signals (PC)
to the engine actuators and to ensure adequate vehicle stability by compensating the guidance and control loop with proper attitude e r r o r and r a t e filters.
These f i l t e r s a r e implemented with passive elements.
The FCC shown in Slide 6 has as inputs attitude-error steering commands ( A4's) from the LVDA and body r a t e s (4's) f r o m the body rate gyros.
These inputs, when filtered in the FCC and summed, result in actuator commands (PC's)which move the engine actuator, and thus the thrust vector, to
cause changes in vehicle dynamics.
�2. FCC Simulation Model
Slide 8 shows a simplified block diagram of the pitch control loop,
including sloshing and bending models for a single engine stage. The simulated attitude and attitude rate filters a r e implemented with passive elements
a s in the actual FCC. The control gains a. and a 1 a r e changed during actual
and simulated flight in discrete steps to offset changes in the control moment
of inertia.
D. Engine Actuators
1. Actual Actuators
The Saturn engine actuators, while differing in type from stage-to-stage,
a r e all highly nonlinear with rate and position limits.
2 . Actuator Simulation Models
Linear approximations of the engine actuators a r e used in the hybrid
simulation. The transfer functions for these actuator models a r e of order
three o r four, depending on the boost stage. In using these linear approximations, small engine angles (0 to 1 3 degrees) a r e assumed. Simplified
nonlinear actuator models a r e presently being developed to handle more
severe cases of engine movement.
E. Bodv Bending Model
The effects of one mode of body bending, caused by forces due to engine
position ( p ) and acceleration
a r e simulated in the pitch plane (Slide 8) by
a second order linear model. The effects of bending (dB and A O B ) a r e sensed
by the rate gyros and platform gimbals and therefore affect the guidance and
control loop.
(B),
F. Moment Equation Model
Motion of the rigid body is described by simple rotational mechanics:
is equal to a moment coefficient C2 times Sin P ,
The attitude acceleration (4)
where C2 is total thrust times the moment a r m (distance from engine gimbal
to center of gravity) divided by the moment of inertia. The Hybrid Simulation
�computes C2 on line from its component parts. The small angle approximation Sin P = P (radians) is used.
G. Propellant Slosh Model
Propellant Sloshing (LOX and LH2) effects on the vehicle attitude acceleration in the pitch plane and the Propellant Utilization System Valve Control
a r e included in the simulation of the second and third stages. The inputs that
cause major sloshing action in the pitch plane a r e the vehicle translational
acceleration due to thrust in the pitch plane and vehicle attitude acceleration
in the pitch plane. These inputs cause the propellants to move against the
tank walls which, in turn, causes attitude acceleration to be induced by two
factors: the force of the propellant sloshing m a s s acting on tank walls through
a moment a r m about the center of gravity, and the sloshing m a s s being displaced from the center line of the vehicle acting through a radial moment a r m
about the center of gravity. The model used to simulate the sloshing effect
during second and third stages consists of two linear second-order directly
coupled differential equations with m a s s varying coefficients. These differential equations describe the radial motion of the slosh m a s s (LOX and LH2)
in the pitch plane.
H. S-I1 and S-TVB Propellant Utilization Systems (PU)
1. Actual PU System
One function of the PU system is to control the LOX and LH2 flow to the
thrust chamber in such a manner that depletion of LH2 and LOX will occur
simultaneously. Remaining LOX and LH2 a r e measured by capacitance-type
probes in each tank. The signal from each probe is gain adjusted so that when
the two signals a r e differenced, a resulting signal will drive the PU valve
position servo to give a desired EMR.
Sloshing (LH2 and LOX) causes the signals representing remaining propellant to vary, which in turn tends to cause the PU valve position, EMR, and
thrust to vary at the sloshing (LOX and L H ~ frequencies.
)
Since a varying
thrust has ill effects on the guidance system, the PU valve control signal is
filtered so that sloshing frequencies a r e highly attenuated in the resulting
valve control signal.
�2. PU Simulation Models
Slide 9 depicts a S-IVB PU system model obtained from a MSFC Guidanct
This is basically the model used in the Hybrid
Dynamics Design Document.
Simulation. The S-I1 model is essentially the same except for a different
sloshing filter. In the PU model, remaining LH2 and LOX masses at any point
in time a r e determined by integrating the total flow r a t e s and subtracting these
f r o m the initial LH2 and LOX masses. Slide 9 also shows how the effects of
propellant sloshing on the PU system are implemented.
VI. COMPUTING TASK ASSIGNMENTS
In assigning computing tasks several factors were considered that seemed
to fall into two general categories as shown in Slide 10 and as follows:
A. Application Orientation
1. Frequency Content
In the Hybrid Simulation, models with high frequency content were simulated on the analog computer, which has a bandwidth of several kH. The f r e quency response of digital computers depends on both the algorithms used to
represent a given model and the solution r a t e of the algorithms. In general,
f o r good accuracy, the solution rate must be several times the highest significant frequency .
2. Precision Requirements
Where high precision was required the digital computer was used, because parameter value can be maintained and expressed in much smaller
increments. An analog signal value is usually expressed in no more than
four o r five decimal places.
3 . System Composition
In the r e a l world the LVDC is digital while the FCC is analog. Precision
in a simulation need not be greater than it is in the actual system. This was
a consideration in assigning the FCC simulation task.
MSFC Memo #R-ASTR-F-66-45, Phase I1 Guidance Dynamics Design Document f o r AS-207, 8 March 1966.
�Flight Hardware Interface
4.
The output of the hardware (LVDC-LVDA) consists of attitude steering
commands ( A 4 ) which a r e analog signals. The inputs to the flight hardware
a r e platform velocity counts, which a r e discrete in nature, and platform gimbal angles, which a r e analog signals.
5. Type of System
One important consideration in some of the model assignments was the
time varying coefficients in the differential equations representing the models.
The analog computer readily lends itself to modeling the differential equation
while the digital computer easily handles time varying coefficients. The
models were therefore simulated on a hybrid system using multiplying DAC's
to combine the coefficients with the dependent variable and i t s derivatives.
B
.
Simulation Hardware
1. Memory Size and Speed of Digital Computers
While memory size can be important, it was not a consideration in this
simulation. The speed of the digital computer was a consideration, in that i t
determined how fast the various loops in the program could be processed and,
therefore, what the solution r a t e s f o r the various model algorithms would be.
2. Equipment Configuration of the Analog Computer
Task assignment is largely dependent on the types and number of elements on the analog computer. In this simulation the analog computer specifications were based on the already determined task assignments listed below;
thus the elements were not really a consideration in this case.
3 . Linkage Characteristics
The bit configuration (word length), conversion rates, and number of
channels @/A, A/D, D/D) were important considerations in assigning c e r tain tasks.
�4. Communication Channel Count and Precision Versus Consolidation of
Small Computing Task on One Machine
When implemented all-analog o r all-digital by the use of special techniques, a model which is best suited to hybrid application will use fewer conversion channels but may lose precision and accuracy.
C
.
Assignments
Slide 11 shows a list of the actual computing task assignments.
a r e as follows:
These
Analog
-
-
-
Flight Control Computer
Control Actuator
Moment Equations
Propellant Slosh Model
Body Bending Model
Propellant Utilization Control System
Digital
-
-
D.
Inertial Platform
Time and Mass Varying Function Generation
Navigation Model
Telemetry Ground Station
Control of Automated Setup and Checkout of Analog Computer
Propellant Utilization System Valve and Pump Model
Data Reduction and Preparation
Hybrid Simulation Tie -In
Tie-in of the total hybrid simulation is presented by tracing system
signal flow in the S-IVB stage pitch plane as follows. The analog computer
receives the attitude steering command ( A 4 ) from the LVDC-LVDA. The
P
control computer model filters this signal and sums it with the filtered body
attitude rate ($p) generated by the vehicle dynamics model. The resulting
signal (PCP)is fed to the actuator model which generates the simulated engine
�(9).
angle
This is then multiplied by the control moment coefficient (CZp),
provided by a digital model, to result in a rigid body pitch attitude acceleration
due to engine thrust. This component is summed with the attitude acceleration
arising from propellant sloshing to give total attitude acceleration ($p). This
$p is integrated once to result in rigid body attitude rate (dp)that is summed
with attitude rate due to body bending
) to give a simulated body rate gyro
bp
output (bgp). The propellant sloshing model has a s inputs, engine angle ( 6 )
P
and attitude acceleration (Pp), while the bending model is forced by engine
angle (Pp) and engine rotational acceleration (b ).
P
(6
The propellant utilization (PU) system is forced by remaining LOX and
LH2 m a s s which a r e computed on a digital computer. The PU system is
perturbed by radial positions of propellant sloshing m a s s e s (LOX and LH2).
The output of the PU system is valve position (av).
A digital computer receives as inputs from the analog models the
attitude rate (dgp), the engine angle (Pp), and the PU valve position (8,). The
$gp is resolved into the inertial platform frame to result in a platform gimbal
angle rate ( 6 p), which is integrated to give the gimbal angle (0 p). The PU
valve position (6 ). is used in a table look scheme of total thrust ( F ) versus
6, to obtain F, which is then divided by remaining vehicle m a s s (mv) to result
in total vehicle acceleration (F/mv). Vehicle m a s s is computed by subtracting
integrated propellant flow rates from initial vehicle mass. Components of body
m ~ the engine gimbal angles.
acceleration a r e obtained by resolving ~ / through
These components a r e in turn resolved through the inertial platform gimbal
angles to result in inertial platform accelerations components, which a r e used
to compute changes in measured platform velocities. These changes in velocities
and the simulated inertial platform gimbal angles a r e transmitted via special
interface equipment to the LVDC -LVDA where the accumulated velocity changes
a r e used for navigation and computation of commanded platform gimbal angles
(x). These x ' s when differenced with the actual platform gimbal angles result
in attitude e r r o r signals (+) which a r e transformed into the body frame to
result in attitude steering commands (A$). These A $ 's a r e then transmitted t o
the analog computer to close the guidance and control loop.
�VII. RESULTS
In concluding this presentation, some of the results from the AS-204
L M / ~Hybrid Simulation studies will be discussed. Slides 12 and 13 show
s t r i p recordings of some of the simulated vehicle and systems dynamics f o r
a nominal case. From left to right on Slide 12 a r e the pitch attitude steering
command (A$*), pitch engine angles (Bp (1,2) and ($ (3,4), pitch body r a t e
(&p), yaw attitude steering command ( ~y),
4 yaw engine angles (fly (2,3) and
(Py (1,4), and yaw body rate (&y). Magnitudes of these parameters and significant flight events a r e indicated. Slide 13 shows Cfrom left to right) the
roll attitude command @OR), body roll r a t e (eR), radial displacement of the
fuel and LOX sloshing m a s s (Z and ZL), remaining LOX and fuel weight
(WL and WF), the change in engine mixture ratio (AEMR), and P U valve position (aV). These simulation results compare favorably with actual flight data.
All significant differences a r e attributable to uncertainties in certain initial
conditions. While results from the hybrid simulation have been good, certain
anomalies in actual Saturn flights have revealed a need f o r studies that were
not previously considered. These studies can be handled by making slight
modifications to certain simulation models.
VIII.
SUMMARY
A hybrid simulation is used to dynamically verify the guidance and cont r o l subsystems of Uprated Saturn I and Saturn V vehicles. This is done by
simulating pertinent vehicle dynamics in a closed guidance and flight control
loop with flight-type (LVDC and LVDA) hardware in the loop. In conjunction
with equipment constraints, the computational complexity imposed by the
requirements for high precision solutions of the guidance and navigation equ:
throughout the total boost portion of the mission and f o r accurate solution of
nonlinear differential equations with variable coefficients and high frequenc:
content led to the choice of a hybrid system for this application. This simr
has already proven to be a valuable tool fop preflight prediction of vehicle/
system dynamic performance and i t s effects on guidance accuracy. With
further developments, it will also be useful for detailed evaluation of dyna.
behavior under extreme combinations of off -nominal situations.
�LAUNCH ESCAPE
SYSTEM
LAUNCH ESCAPE
SYSTEM
COMMAND
MODULE
SERVICE
MODULE
COMMAND
MODULE
SERVICE
MODULE
ADAPTER
S-IVBSTAGE
LUNAR
LE
q+- - -
I
1 J -2 ENGINE
INSTRUMENT
S-IVB STAGE
5-IVB AFT
INTERSTAGE
--I I
. 1 J-2 ENGINE
S -IC STAGE
S-IB STAGE
8 H-1 ENGINES
APOLLO-SATURN V VEHICLE
UPRATED SATURN I VEHICLE
Slide 1
�SATURN INSTRUMENT UNIT
Slide 2
�Instrument Unit Subsystems
e
STRUC'TURAL PORTION
e
GU IDANCE AND CONTROL
e
ENV I RONMENTAL CONTROL
e
MEAS UR ING AND TELEMETRY
e
RAD 10 FREQUENCY
e
ELECTR ICAL
Slide 3
15
�Guidance and Control Subsystem
o
LAUNCH VEHICLE DIGITAL COMPUTER (LVDC)
o
LAUNCH VEHICLE DATA ADAPTER (LVDA)
o
FLIGHT CONTROL COMPUTER (FCC
o
ST 124 INERTIAL PLATFORM
a
BODY RATE GYROS
Slide 4
16
�Simplified Saturn V
Slide 5
17
�Guidance and Control System
- ..
..
..
Xm ' Ym ' Zm
lnertial
Platform
Assernbl y
>P
-------------------
B~
-
I
I
I
-%
JJ
I
I
inertial Velocities (~11,A V , OW)
Inertial Platform Gimbal Angles ( 0 )
1
__c
LVDC
-1
-1
LVDA
n v ~
b
Flight
Control
Computer
@CP
Engine
Actuator
Vehicle
.
-@%
Dynamics
I
A
A '
-vT'
'
PU
System
Rate
Gyro
---------
Slide 6
18
I
I
I
�FLIGHT COMPUTER TASKS B Y PHASE
185 Kilometers (100 Naut. Miles
I
SATURN V I APOLLO
PHASE
t
Slide 7
19
I
TASK
I
�Control Loop
Slide 8
�Propellant Utilization System
z~
,c o s e F
SLOSH
INITIAL
L H 2 ~ t ~LH2~ ~ ~
--
.
8wF
a z ~
FROM SLOSH MODEL
NOMINAL L H ~
FLOW RATE
I
LH2 BRIDGE SERVO
I
'PU
K~ ( d s 2
cs5
t
6s4
-
- bS '
-
es3 fs2 g s
POSlTlONER
VALVE
AA\P
FORWARD SHAPING
-1
VALVE
:
K~
I
GEARS & POT
REDBACK
I
LOX BRIDGE SERVO
-
-
2C~"~
,
1
1\ .s
K~~
z~
FROM SLOSH MODEL
-
'
coseL .
IINITIAL
aW~
aZL
LOX SLOSH
WEIGHT
Slide 9
LOX
WEIGHT
I
N O M l NAL LOX
FLOW RATE
Lox
TO
THRUST
MODEL
�Factors Related to Computing Task
Assignments in Hybrid Simulation
1. APPLICATION
a.
b.
c.
d.
e.
2.
FREQUENCY CONTENT
PREC IS ION REQU I REMENTS
SYSTEM COMPOS IT1ON
FLIGHTEQUIPMENT INTERFACE
TYPE OF SYSTEM
COMPUTER HARDWARE
a.
b.
c.
d.
MEMORY S I Z E AND SPEED OF D I G l T A L MEMBER
EQU I PMENT CONFIGURATION OF ANALOG MEMBER
LINKAGE CHARACTER1 STICS
COMMUNICATION CHANNEL COUNT AND PRECISION
vS
CONSOLIDATION OF SMALL COMPUTING TASK ON ONE MACHINE
Slide 10
22
�Computing Task Assignment
ANALOG
a
a
a
a
a
a
FLIGHT CONTROL COMPUTER
CONTROL ACTUATOR
MOMENT EQUATIONS
PROPELLANT SLOSH MODEL
BODYBENDINGMODEL
PROPELLANT UTl L l ZATl ON CONTROL SYSTEM
DIGITAL
a
a
a
a
a
0
0
lNERTl AL PLATFORM
TIME ANDIOR MASS VARYING FUNCTIONS GENERATION
NAVIGATION MODEL
TELEMETRY GROUND STATION
CONTROL OF AUTOMATED SET-UP AND CHECKOUT OF ANALOG
PROPELLANT UTILIZATION SYSTEM VALVE AND PUMP MODELS
DATA REDUCTION AND PREPARATION
Slide 11
23
�Slide 12
�Slide 13
�4
Federal Systems Division, Space Systems Center, ~untsville,Alabama
�
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Saturn V Collection
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<a href="http://libarchstor.uah.edu:8081/repositories/2/resources/60" target="_blank" rel="noreferrer noopener">View the Saturn V Collection finding aid in ArchivesSpace</a>
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Saturn V Collection
Description
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<p>The Saturn V was a three-stage launch vehicle and the rocket that put man on the moon. (Detailed information about the Saturn V's three stages may be found<span> </span><a href="https://www.nasa.gov/centers/johnson/rocketpark/saturn_v_first_stage.html">here,<span> </span></a><a href="https://www.nasa.gov/centers/johnson/rocketpark/saturn_v_second_stage.html">here,<span> </span></a>and<span> </span><a href="https://www.nasa.gov/centers/johnson/rocketpark/saturn_v_third_stage.html">here.</a>) Wernher von Braun led the Saturn V team, serving as chief architect for the rocket.</p>
<p>Perhaps the Saturn V’s greatest claim to fame is the Apollo Program, specifically Apollo 11. Several manned and unmanned missions that tested the rocket preceded the Apollo 11 launch. Apollo 11 was the United States’ ultimate victory in the space race with the Soviet Union; the spacecraft successfully landed on the moon, and its crew members were the first men in history to set foot on Earth’s rocky satellite.</p>
<p>A Saturn V rocket also put Skylab into orbit in 1973. A total of 15 Saturn Vs were built, but only 13 of those were used.</p>
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spc_stnv_000055
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"A Hybrid Simulation for Dynamic Verification of Saturn Guidance and Control Subsystems."
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IBM No. 68-U60-0013
Description
An account of the resource
This paper presents a discussion of a hybrid simulation used to dynamically verify the Saturn Guidance and Control subsystems. First, the Saturn vehicle is briefly described to provide background information. The Instrument Unit (IU) is considered in more detail to give a proper setting for the Guidance and Flight Control (G and FC) discussion that follows. After a brief description of the actual G and FC System operation, simulation models of the G and FC components are considered in detail. This is followed by a discussion of the model assignment to a particular computer (digital or analog) and justification for making that assignment. Finally, results of the AS-204/LM1 hybrid simulation studies are briefly considered with mention of the actual flight data.
Creator
An entity primarily responsible for making the resource
Patray, Ronald T.
International Business Machines Corporation. Federal Systems Division
Date
A point or period of time associated with an event in the lifecycle of the resource
1965-07
Temporal Coverage
Temporal characteristics of the resource.
1960-1969
Subject
The topic of the resource
Saturn Project (U.S.)
Project Apollo (U.S.)
Saturn launch vehicles
Apollo spacecraft
Spacecraft guidance
Spacecraft control
Computerized simulation
Dynamic tests
Type
The nature or genre of the resource
Reports
Text
Source
A related resource from which the described resource is derived
Saturn V Collection
Box 26, Folder 34
University of Alabama in Huntsville Archives, Special Collections, and Digital Initiatives, Huntsville, Alabama
Language
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en
Rights
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This material may be protected under U. S. Copyright Law (Title 17, U.S. Code) which governs the making of photocopies or reproductions of copyrighted materials. You may use the digitized material for private study, scholarship, or research. Though the University of Alabama in Huntsville Archives and Special Collections has physical ownership of the material in its collections, in some cases we may not own the copyright to the material. It is the patron's obligation to determine and satisfy copyright restrictions when publishing or otherwise distributing materials found in our collections.
Relation
A related resource
spc_stnv_000051_000074