Computer programs provide preliminary trajectory and guidance information required for feasibility studies in space mission analysis. The advanced mission analysis computer programs include programs for approximate solutions, programs for targeting and output, and programs for Monte Carlo and linear guidance analysis

Bjorkman, W. S.

Brooks, M. J.

English

NASA Technical Reports Server

June 1969	 Brief 69-10171 
NASA TECH BRIEF	
IL Ali 
NASA Tech Briefs are issued to summarize specific innovations derived from the U.S. space program, 
to encourage their commercial application. Copies are available to the public at 15 cents each 
from the Clearinghouse for Federal Scientific and Technical Information, Springfield, Virginia 22151. 
Advanced Mission Analysis Programs 
The problem: 
Space mission analysis requires consideration of 
numerous parameters which affect mission success. In 
particular, a suitable trajectory must always be found 
which satisfies imposed mission constraints. 
The solution: 
A group of computer programs designed to provide 
preliminary trajectory and guidance information re-
quired for feasibility studies in mission analysis. 
How it's done: 
The Advanced Mission Analysis Computer Pro-
grams fall into three categories. 
I. Programs for approximate solutiops 
2. Programs for detailed targeting and output 
3. Program for Monte Carlo and linear guidance 
analysis. 
The first category includes all those programs 
whose solutions are based on single-conic-section ap-
pr@ximations for trajectory definition. These solutions 
may be used as acceptable estimates of the exact solu-
tions for many problems or as "first-cut" iteratives for 
detailed targeting. The second category includes the 
programs for detailed targeting and for generating 
detailed output along the trajectory. These programs 
"patch" conic sections when the trajectory changes 
gravitational centers to provide more realistic trajec-
tory simulations than the programs of category 1. The 
program of category 3 also uses a patched-conic tra-
jectory simulation model in the process of studying 
the effects of injection errors and midcourse guidance 
corrections on terminal conditions. 
The programs in category I are further categorized 
according to the mission(s) each one treats. The mis-
sions which may be treated are: 
a. Planet-to-planet transfers (PLANET) 
b. Moon-to-Earth transfers (MOON) 
c. Earth-to-Moon transfers (MOON, QUIMP) 
d. Circumlunar gravity-assist (CIRCUM) 
e. Interplanetary gravity-assist (GRASS) 
f. Launch-to-orbit rendezvous (RENDEZ) 
g. Low-thrust interplanetary rendezvous (LTOPT) 
The approximate solution obtained for most of 
these missions includes the computer values for a set 
of control variables connecting the transfer orbit to 
a parking orbit at the launch body. When the transfer 
trajectory is an escape trajectory from the launch body 
(a, b, c above) this computation is done in subroutine 
LAUNCH. When Earth-to-Moon trajectories are 
being considered, the connection between the parking 
orbit and the transfer orbit is a fundamental part of 
the solution for the transfer orbit as computed by 
MOON or QUIMP. 
A capability for parametric variation of launch date 
arid flight time is provided with the programs for ap-
proximate solutions. These two parameters comprise 
the basic input for any single case, so that by asking 
for their variation, many cases may be automatically 
generated without additional input. In addition to this 
feature, the trajectory and parking orbit characteristics 
associated with each launch dateflight time pair may 
be systematically stored on tape for later printing or 
plotting. 
The programs in category 2 are intended to provide 
more detailed answers for the mission analyst than he 
could obtain from the approximate program. The two 
general capabilities included here are: 
a. Targeting (REFINE) 
b. Detailed output (OUTPUT)
(continued overleaf) 
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Both of these capabilities make use of the patched-
conic scheme for trajectory simulation. This scheme, 
while still only an approximation to an exact simula-
tion treating all known perturbations, has proven ac-
ceptably valid and relatively fast for preliminary an-
alysis of most space missions. 
Under the patched-conic assumption, the trajectory 
is considered to be a Keplerian conic section at any 
given time. When the vehicle reaches the "sphere of 
influence" of its gravitational center or central body, 
however, the conic section upon which it is moving is 
changed to one relative to the primary (Sun or Earth). 
Or, when the sphere of influence of another massive 
body (Moon or planet) is reached while the conic is 
focused on the primary, the conic is changed to be 
relative to that body. The spheres of influence are 
considered to be constant in radius. 
The guidance analysis program is a tool for study-
ing the effects of injection and midcourse guidance 
system errors on the achievement of mission objec-
tives for lunar and interplanetary missions. The 
patched-conic trajectory package previously men-
tioned is used for all of the trajectory calculations re-
quired in the guidance analysis. The program is ca-
pable of both linear (propagation and midcourse ad-
justment of the covariance matrix of state errors) and 
Monte Carlo (sampling of random errors) treatments. 
As many as three midcourse guidance correction 
maneuvers may be simulated for any mission. Each 
maneuver is formulated as a velocity impulse to be 
applied at a specific time. The guidance laws which 
are available are: 
a. Fixed time of arrival, attitude unconstrained 
b. Fixed energy at arrival, attitude unconstrained 
c. Minimum fuel, attitude unconstrained 
d. Miss vector control, ecliptic plane-constrained 
e. Minimum fuel, ecliptic plane-constrained
f. Earth-pointing 
g. Sun-pointing 
h. Target body-pointing 
i. Injection velocity-pointing 
j. Star-pointing 
k. Normal to the ecliptic plane 
Each of the three maneuvers for any mission may 
be based on a different guidance law. Errors in execu-
tion of each midcourse maneuver are considered for 
both the linear and Monte Carlo treatments. The ex-
ecution error types are: 
a. Proportional to the commanded correction 
b. Resolution or bias 
c. Pointing of the commanded correction 
Input of the initial state and the initial state error 
covariance matrix may be assumed at insertion, im-
mediately prior to the injection thrusting period, or at 
injection. The program will update the state and the 
covariance matrix to injection, if necessary, including, 
linearly, the effects of injection thrust errors. 
Notes: 
I. These programs are written in FORTRAN IV 
and MAP for use on the IBM 7094 computer. 
2. Inquiries concerning these programs should be 
directed to:
COSMIC 
Computer Center 
University of Georgia 
Athens, Georgia 30601 
Reference: B69-10171 
Patent status: 
No patent action is contemplated by NASA. 
Source: W. S. Bjorkman and M. J. Brooks 
of Philco-Ford Corporation 
under contract to 
Goddard Space Flight Center 
(GSC-10575)
S 
S 
. 
Brief 69-10171	 Category 06


Advanced mission analysis programs

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