LANDSAT-5 ORBIT ADJUST MANEUVER REPORT 

Prepad for 

NATIONAL AEROMACITICS AND SPACE ADMINISTRATION 

Goddad Space Flight Center 

Greenbelt, Maryland 

CONTRACT NAS 5-27888 

Task Assignment 14300 

JUNE t9&4 

I(P84-10155) LAhDShT-5 OBBIT~DJOST LLIlEOVEB 

, BEPOBT (Computer Sciences Cor~,) 63 p 

N84-29276 1 

CSC 

COMPUTER SCIENCES CORPORATION 

-- - .I - * - - . * t-3 .r . 
LANDSAT-5 ORBIT ADJUST 

MANEUVER mPORT 

Prepared for 

GODDARD SPACE FLIGHT CENTER 

BY 

COMPUTER SCIENCES CORPORATION 

Under 

Contract NAS 5-27888 

Task Assignment 14900 

Prepared by: Approved by: 

??A I. 

P. 3. ~&sett Date Date 

Wical Supervisor 

R. L. ~ohnsdnu Date 

~uhctional Area Manager 
ABSTRACT 

The Landsat-5 spacecraft was successfully launched from the 

Western Test Range by a Delta 3920 launch vehicle on 

March 1, 1984. This document describes the orbit adjust 

manewers performed to raise the spacecraft to mission altitude, synchronize it with the required groundtrack, and 

properly phase the spacecraft with Landsat-4 to provide an 

8-day full Earth coverage cycle. It also describes maneuver 

planning and evaluation procedures, data and analysis results for all maneuvers performed to date, the frozen orbit 

concept, and the phasing requirement between Landsat-4 -nd 

Landsat-5. 
TABLE OF CONTENTS 

Section 1 - Introduction. . . . . . . . . . . . . . . . 1-1 

Section 2 - Maneuver Planning and Evaluation. 

2.1 Landsat-5 Mission Orbit ~equirements . . . . . . . 2-1 

2.2 Frozen Orbit Concept . . . . . . . . . . . . . . . 2-1 

2.3 Phasing with Landsat-4 . . . . . . . . . . . . . . 2-5 

2.4 Software for Maneuver Planning and Evaluation. . . 2-7 

2.5 Premaneuver Planning . . . . . . . . . . . . . . . 2-8 

2.6 Postmaneuver Evaluation. . . . . . . . . . . . . . 2-9 

- Section 3 - Landsat-5 Maneuvers . . . . . . . . . . . . 3-1 

3.1 Postlaunch Injection Error Removal Sequence. . . . 3-1 

3.2 Orbit and Groundtrack Xaintenance Maneuvers. . . . 3-9 

wendix - Landsat-5 Orbit Adjust Maneuver Data 

References 

iii 

SECTION 1 - INTRODUCTION 

The Landsat-5 spacecraft was successfully launched from the 

Western Test Range by a Delta 3920 launch venicle on 

March 1, 1984. The launch was very near nominal and resulted in an orbit that was approximately 12 kilometers below the final mission altitude. The orbit was targeted low 

intentionally to ensure that no orbit lowering maneuvers 

would be required (this would necessitate a 180-degree yaw 

of the spacecraft, which is undesirable). A series of eight 

orbit raising maneuvers was performed between March 7 and 

April 4, 1984, to raise the semimajor axis the remaining 

12 kilometers. The maneuvers were performed at the proper 

times so that both phasing with the World Reference System 

(WRS) groundtrack grid and an 8-day coverage cycle between 

Landsat-4 and Landsat-5 were achieved. The series of maneuvers also achieved a frozen orbit. Periodic orbit maintenance maneuvers have kept the groundtrack within the 

required bounds. 

This document follows the format of the Landsat-4 Orbit At?- 

just Maneuver Report (Reference 1) prepared by 

R. J. McIntosh. Section 2 defines the Landsat-5 orbit requirements, discusses computer software and operational procedures used for maneuver planning and evaluation, and 

briefly describes the frozen orbit concept. Also discussed 

is the phasing requirement between Landsat-4 and Landsat-5 

needed to achieve the 8-day coverage cycle. Section 3 describes the postlaunch injection error removal maneuver sequence and discusses orbit maintenance maneuvers. The 

appendix contains data and analysis results covering all 

maneuvers performed to date. It is intended that update 

pages will be published for insertion into the appendix as 

future maneuvers ars performed. 
SECTION 2 - MANEUVER PLANNING AND EVALUATION 

This section defines the Landsat-5 mission orbit and groundtrack and discusses computer software and procedures used 

for premaneuver planning and postmaneuver evaluation. In 

addition, brief descriptions of the frozen orbit concept and 

the phasing requirement between Landsat-4 and Landsat-5 are 

given. 

2.1 LANDSAT-5 MISSTON ORBIT REQUIREMENTS 

The nature of the Landsat-5 missio.~ requires that the spacecraft orbit have the proper altitude and inclination to 

maintain a Sun-synchronous node rate (mean local time at any 

descending node crossing is constant) and a 16-day groundtrack repeat cycle. The number of orbits in the groundtrack 

repeat cycle is 233; that is, after 233 orbits, the spacecraft must cross over the same longitude point on the 

4. Earth's Equator. The nominal mean semimajor axis required 

for this repeat cycle is approximately 7077.8 kilometers. 

The WRS groundtrack grid defines a series of descending node 

crossings equally spaced around the Earth's Equator, approximately 172 kilometers apart, with the base longitude 

defined as 295.4 degrees east longitude. An orbital inclination of approximately 98.2 degrees maintains the required 

Sun-synchronous nodal regression rate and mean local time of 

the descending node (between 9:30 and 10:OO a.m.). The or- 

 it requirements are defined in Reference 2. 

2.2 FROZEN ORBIT CONCEPT 

Low-altitude circular orbits are subject to strong perturbations from the oblate Earth's gravitational potential. The 

magnitude of the effects of these perturbations depends on 

the initial values of certain orbital parameters, namely, 

semimajor axis, eccen-ricity, inclination, and argument of 
perigee. By targeting toward the proper values of these 

orbital parameters, the effects of certain perturbations can 

be minimized. This is the case of the frozen orbit, in 

5: . +- which the line of apsides (the line joining apogee and perigee) is stopped or frozen. For a near-circular orbit, the 

changes in the average argument of perigee (3) and the 

average eccentricity (g) become zero when w equals 90 aegrees and e approaches some determinable small value. This 

value of Z (the frozen eccentricity) is a function of the 

averaged semimajor axis (a) and inclination (i) . The frozen 

eccentricity for the Landsat-5 orbit is approximately 

0,0012. For initial values of G near 90 degrees and initial values of near the frozen eccentricity, the averaged 

argument of perigee and eccentricity oscillate within a 

small range about the frozen condition. Figure 2-1 shows 

the evolution of G and G with time for a near-frozen 

Landsat-5 orbit. 

The evolution of a near-frozen orbit can be described in an 

eccentricity vector space, where the motion of the eccentricity and argument of perigee about the frozen point is 

circular. An example of a satellite with E and w near the 

frozen values is illustrated in Figure 2-2, using the eccentricity vector space representation (e cos G and e sin 

W). Any set of initial conditions, such as point R in 

Figure 2-2, results in the cyclic evolution of the 5 cos 

and 5 sin parameters. The frozen orbit concept is developed in Reference 3. 

The main advantage of the frozen orbit is a minimum variation in altitude above any given latitude due to a near- 

?, constant perigee location. A near-constant altitude above a 

given latitude will minimize the necessary geometric corrections to the Landsat-5 images. The frozen orbit is a derived requirement for Landsat-5. The eccentricity will 
0**4 '1 X TY C'I I ECCENTE MEAN 

<; therefore be maintained well below 0.003 for nominal image 

correction (Reference 2). By performing the postlaunch maneuvers (required to achieve mission orbit altitude and 

groundtrack phasing) at the optimum location, a near-frozen 

orbit can be reached. Once the initial target frozen orbit 

i~ reached, Z and i5 can be controlled to some extent as 

part of the routine orbit maintenance maneuvers (which are 

required to counteract the effects of atmospheric drag) 

without additional fuel requirements. Targeting toward the 

frozen orbit was done during the postlaunch injection error 

removal maneuver sequence for Landsat-5, as discussed in 

Section 3. 

- PHASING WITH LANDSAT-4 

Landsat-4 and Landsat-5 both Pave a repeating groundtrack 

every 233 revolutions (16 days,. Figure 2-3 represents a 

one-revolution segment of the Equator that contains 16 WRS 

intervals. Each dot on the Equator represents a WRS longi- 1' 

tude point. The number over each WRS longitude shows the 

day in the 16-day repeat cycle when a Landsat spacecraft in 

a nominal mission orbit would cross (from north to so'uth) 

the WRS point. Day 0 of the 16-day repeat cycle was arbitrarily chosen. 

The nominal mission orbits of Landsat-4 and Landsat-5 are 

identical except for tk.eir phasing. Landsat-5 is requi~ed 

to be phased with Landsat-4 such that full Earth coverage is 

provided every 8 days. The 8-day full Earth coverage cycle 

is achieved when the Landsat spacecraft are 180 degrees 

apart when in the same orbital plane, as shown in Figure 2-4. This phasing configuration also minimizes interference between the two satellites. Figure 2-5 contrasts 

the days of coverage in the 16-day cycle for a segment or 

the Equator for Landsat-4 and Landsat-5. The figure depicts 

Landsat-5 as being correctly phased eight WRS intervals from 

6 Landsat-4, thus providing full coverage every 8 days. 
ORIGINAL PAGE I$ 

OF POOR QUALITY 

0,16 9 2 11 4 13 6 15 8 1 10 3 12 5 14 7 0.16 - - - - - - - - - - - - - - - - - - 

I+ 1 REVOLUTION 4 

Figure 2-3. WRS Longitude Points and the 16-Day Coverage Cycle 

Figure 2-4. Proper Landsat-5 Phasing With Landsat-4 for 

Full Earth Coverage Every 8 Days 

8 1 10 3 12 5 14 7 0,16 9 '2 11 4 13 6 15 8 LANDSAT-5 

1 REVOLUTION ~-4 9667184 

Figure 2-5. Tight-Day Coverage Cycle Between Landsat-4 

and Landsat-5 
2.4 SOFTWARE FOR MANEUVER PLANNING AND EVALUATION 

The software used for Landsat-5 maneuver support consists of 

programs to perform high-precision orbit propagation, 

groundtrack monitoring, maneuver targeting, propulsion system modeling, and tracking station coverage prediction. 

These programs are as follows: 

a Goddard Mission Analysis S:.stem (GMAS) --The GMAS 

Cowell propagator is used to generace high-precision ephemeris (EPHEM) files for use by the cther programs. .The GMAS 

TRACK module is used to cietermine groundtrack errors. Averaged orbital elements are generated by the AVECOX utility 

module. 

General Maneuver Program (GMAN)--GMAN performs maneuver targeting, maneuver reconstruction, and propulsion 

system modeling. 

a Groundtrack Monitoring Prcgram (GNDTRAK)--GNDTRAK 

reads a standard EPHEM file, interpolates to find descending 

node crossings, and compares the node crossings to the required groundtrack grid. rile mean local time of each crossing is also output. 

a Circular Orbit Restoration Program (RESTOR)--RESTOR 

determines maneuver requirements to achieve target values of 

mean semiiaajor axis, eccentricity, and argument of perigee. 

RESTOR was developed for frozen orbit targeting. 

a Acquisition Data Program (ACQSCAN)--ACQSCAN determines acquisition- and loss-of-signal (AOS and LOS) times 

for selected ground stations, determines shadow times, generates various reports, and provides coverage schedules for 

several Tracking and Data Relay Satellites. ACQSCAN reads a 

standard EPHEM f i1.e. 

a Ephemeris File Writer Program (EPHGEN)--EPHGEN contains the same Cowell propagator as GMAS; however, EPHGSN 
c, executes in less than 200K bytes of core. The primary output from EPHGEN is a standard EPHEM file. 

2.5 PREMANEUVER PLANhING 

Tnis subsection discusses the procedures generally followed 

in planning a typical orbit adjust maneuver and how the 

software described in the previous subsection is used. The 

following steps are taken in planning a maneuver: 

1. Obtain the latest orbit determination s~lution 

(EPHEM tape) or generate an EPHEM file using GMAS 

or EPHGEN and the latest vector. 

C) 

L. Run GNDTRAK to check the groundtrack error and 

determine when a maneuver is necessary to control 

the groundtrack evolution. 

3. Run the GMAS AVECON utility to generate averaged 

orbital elements at the expected maneuver time. 

4. Input averaged elements to RESTOR to determine maneuver location and magnitude required to control 

orbit semimajor axis, eccentricity, and argument of 

perigee to the desired values. 

5. Ran ACQSCAN to determine the best station coverage 

time for the maneuver (near the location defined by 

RESTOR) . 6. Obtain approval for the requested maneuver date and 

time. 

7. Obtain the latest fuel system temperatures and 

pressures from the Landsat-5 Control Center. 

8. Run GMAN using maneuver magnitude estimates from 

RESTOR and latest temperatures and pressures to 

model the maneuver and predict fuel usage. 
9. Run the GMAS Cowell propagator and TRACK module 

with several solar flux level estimates tc get a 

groundtrack evolution prediction. 

10. If groundtrack prediction is not satisfactory, 

change burn duration and repeat steps 8 and 9 as 

necessary. 

Following this analysis, the burn start time and duration 

are delivered to the control center. 

The greatest effect on the groundtrack evolution is from 

decay of the semimajor axis due to atmospheric drag. Because tne solar flux level (which influences atmospheric 

density at a given altitude) cannot be accurately predicted, 

it is necessary to use several constant solar flux values 

(actually, atmospheric density tables) for predicting the 

effect of a maneuver on the groundtrack. The expected minimum and maximum solar flux levels likely to be encountered 

during some period following the maneuver are used to estimate bounds for the westward groundtrack drift. The objective is to determine an appropriate burn duration so that 

the required groundtrack error limits (510 kilometers) 

will be maintained in case of sudden changes in the solar 

flux. 

2.6 POSTMANEUVER EVALUATION 

Following completion of an orbit adjust maneuver, the following steps are taken to evaluate the propulsion system 

performance ar4 to calibrate the thruster modeling: 

1. Obtain the preburn and postburn orbit determination 

solution EPHEM tapes; retrieve preburn and postburn 

state vectors at the same epoch time (burn end 

time) . 
2. Convert preburn and postburn vectors to averaged 

orbital elements using the GMAS AVECON utility and 

compute the actual change in averaged semimajor 

axis. 

3. Obtain actual temperatures, pressures, and thruster 

durations observed by the Landsat-5 Control Center. 

4. Run GMAN using the observed propulsion system 

parameters to remodel the burn; this generates a 

new prediction of orbital changes. 

5. Convert predicted postburn osculating elements to 

averaged elements (AVECON) and compute predicted 

changes in the orbit. 

6. Compare predicted and observed postburn semimajor 

axis values and compute thruster correction factor, 

7. Perform an attitude thruster burn with GMAN using 

the total observed attitude tnruster pulses; this 

is done to obtain an estimate of fuel used for attitude control. 

Following the completion of this procedure, a postburn analysis report is delivered to the control center. 

The estimated fuel remaining and thrust correction factor 

are cataloged for future use. 

In evaluating the maneuver with GMAN, several assumptions 

are made: 

The attitude is held constant at the value originally commanded by the control center (for example, 

roll, pitch, and yaw = 0.0 degrees). The total 

attitude thruster counts are used to estimate fuel 

usage only. 
The burn time used is equal to the total thruster 

time in milliseconds divided by the number of 

thrusters used (2 or 4). 

For example, GHAN would use a burn time of 251.648 + 4 = 

62.912 seconds with four thrusters firing simultaneously, if 

the observed values for each thruster were as follows: 

Thruster Duration (Milliseconds) 

A1 51,648 

B1 65,344 

C1 67,328 

Dl 67,328 

Total 251,648 (or 251.648 seconds) 

Modeling of the Landsat-5 propulsion system, which is identical to the Landsat-4 propulsion system, is described in 

Reference 4. Since that document was published, an update 

has been made to the equation for thrust in CMAN to model 

thruster warmup. The equations used for modeling thrust and 

specific impulse (ISp) are given below. 

where F = thrust (pounds) 

P = tank pressure (pounds pei square inch absolute 

(psis) 

T = current thruster on-time (seconds; 

TS = steady state time (time for thruster to reach 

full output) 

The last term accounts for thruster warmup. The value of 

TS used in the GMAN spacecraft data file is 20 seconds. 

When T becores greater than TS, GMAN sets the expression 

T/TS equal to 1. 
where ISp = specific impulse (seconds) 

P = tank pressure (psia) 

T'-e thrust calculated by GMAN can be adjusted by adding a 

tlrust correction factor to the program, The thrust is calculated as noted above and then adjusted as follows: 

w lere F = thrust calculated from the polynomial equation 

k = thrust correction factor (nominally = 1.0) 

'ad j = adjusted thrust level 

Tbe thrust correction factor for a maneuver is calculated as 

follows : 

where Aaob, = observed change in averaged semimajor axis 

Aapre = change in averaged semimajor axis predicted by 

GMAN 

kold = thrust correction factor used by GMAN to generate the prediction 

The thrust correction factors for most Landsat-5 maneuvers 

to date have averaged 0.97 (the observed thrust level has 

been 97 ,~trcent of the nominal thrust predicted by GMAN). 
SECTION 3 - LANDSAT-5 MANEUVERS 

The Landsat-5 spacecraft was successfully launched into a 

circular polar orbit on March 1, 1984, by a Delta 3920 

launch vehicle that reached an orbit of 12 kilometers below 

mission ~ltitude. The purpose of the postlaunch maneuver 

sequence was to raise the semimajor axis the remaining 

12 kilometers to the mission altitude, synchronize the 

groundtrack with the WRS grid, obtain a near-frozen orbit, 

and phase the spacecraft with ands sat-4 such that an 8-day 

coverage cycle was obtained between the two satellites. To 

achieve all these goals, maneuver magnitude and timing were 

critical. The following subsections discuss details of the 

postlaunch maneuver sequence and the orbit maintenance n. - 

neuvers. 

3.1 POSTLAUNCH INJECTION ERROR REMOVAL SEQUENCE 

A sequence of eight orbit adjust maneuvers raised the 

Landsat-5 spacecraft to operational altitude. Three short 

burns tested the propulsion system in the primary mode (four 

thrusters) and the bacKup mode (two thrusters). Five large 

maneuvers were performed to raise the orbit and synchronize 

the groundtrack. Details of the maneuvers are presented in 

Table 3-1. A brief description of the maneuvers is given 

below. 

Maneuver 

Number Comments 

1 A 5-second burn with two thrusters to test the 

backup firing mode--This burn produced a semimajor axis change of 197 meters. A large number of attitude thruster firings occurred. 

A 5-second burn with two thrusters to further 

test the backup firing mode--A semimajor axis 

change of 198 meters was achieved. 

A 16-second burn with four thrusters to test 

the primary firing mode--The semimajor axis 

increased by 1.47 kilometers. 
ORIGINAL PAGE 13 

OF POOR QUALITY 
c/ Naneuver 

Number Comments - 

4 A 13-second burn with four thrusters to begin 

the groundtrack phasing sequence--This burn 

produced a change in semimajor axis of 

1.19 kilometers and slowed the groundtrack 

drift rate from 5.8 kilometers per revolution 

to 5.1 kilometers per revolution. The D translational thruster ceased firing during this 

maneuver, causing excessive off-pulsing. The 

onboard timer terminated the burn after 

102 seconds elapsed. The planned burn duration 

for this burn was 51 seconds. 

A 30-second burn performed one orbit after maneuver 4--The burn was designed to complete the 

planned burn duration of 51 seconds started in 

maneuver 4. The burn used the primary firing 

mode, but only three thrusters fired. Thc 

semimajor axis was raised by 2.3 kilometers, 

and the groundtrack drift rate was slowed to 

3.8 kilometers per revolution. 

A 41-second burn with two thrusters--This burn 

raised the semimajor axis 1.9 kilometers and 

slowed the groundtrack drift rate to 2.7 kilometers per revolution. 

A 58-second burn using two thrusters--This burn 

slowed the groundtrack drift rate to 1.1 kilometers per revolution and raised the semimajor 

axis by 2.6 kilometers. 

A 43-second burn with two thrusters-gr his burn 

raised the orbit by 1.9 kilometers, phased the 

spacecraft half an orbital period ahead of 

Landsat-4, and positioned Landsat-5 into a frozen orbit. The drift rate of the groundtrack 

was slowed to 0.03 kilometer per revolution. 

Following maneuver 8, the groundtrack position was approximately 4.6 kilometers west of the required WRS path (well 

~ithin the 210 kilometer limits) and drifting eastward at 

approximately 0.44 kilometer per day. The fuel use estimated by GMAN for these eight burns is 12.6 pounds. 

Table 3-2 presents the predicted and actual changes in averaged semimajor axis for each maneuver and also the calculated thrust correction factors. 
Table 3-2. Predicted and Actual Changes in Averaged Semimajor A

PREDICTED 

POSTBURN B MANEUVER 

1 

2 

OBSERVED PREDICTED I OBSE!VED 

POSTBURN Z PREBURN 

7066.0868 

7066.2883 

THRU

CORRE

FACT

=BOTH PREBURN AND POSTBURN VALUES ARE ONE-ORBIT NUMERICAL AVERAGES AT EPOCH OF BURNOUT FOR EAC

MANEUVER. ALL VALUES ARE IN KILOMETERS. 
Tne evolution of the Landsat-5 groundtrack through the postlaunch injection error removal period is illustrated in 

Figure 3-1. The groundtrack grid comprises 233 equatorial 

crossings that are 172 kilometers apart. At each descending 

node crossing, the groundtrack is compared to the nearest 

grid line. The maximum groundtrack error is then half the 

distance between lines, or 86 kilometers. The initial drift 

rate in the groundtrack following launch was 100 kilometers 

per day. The apparently instantaneous changes from positive 

to negative in the plot of Figure 3-1 indicate that the 

halfway point between two adjacent grid lines has been 

crossed and the groundtrack error is then being checked 

against the next longitude. As maneuvers are performed to 

raise the semimajor axis, the drift rate decreases. The 

final burn was performed when the groundtrack error was 

within the +lo kilometer bounds. 

Figure 3-2 illustrates the evolution of the frozen orbit in 

terms of e cos w and e sin w. The effect of each maneuver on the eccentricity vector can be seen. The appendix 

contains more details on each maneuver. 

Landsat-5 was phased approximately half an orbital period 

(180 degrees) behind Landsat-4 following separation. The 

injection orbit of Landsat-5 was 12 kilometers lower in the 

semimajor axis than that of the Landsat-4 mission orbit. 

This caused Landsat-5 to gradually catch up to Landsat-4 and 

then pass it, until Landsat-5 led Landsat-4 by half an 

orbital period. Orbit adjust maneuver 8 was then executed 

to complete the injection error removal sequence, correctly 

phasing Landsat-5 180 degrees from Landsat-4 to provide an 

8-day full Earth coverage cycle. Figure 3-3 depicts the 

phasing evolution between Landsat-4 and Landsat-5 during the 

injection error removal sequence of Landsat-5. The phasing 

is measured by comparing the times of corresponding 
ORIGINAL PAM 

OF POOR QUALITY 
ORIGINAL PACE I.cl 

OF POOR QUALITY 
ORIGINAL PAGE 19 

OF POOR QUALlW 
descending nodal crossil~gs of the two Landsat spacecrhft. A 

positive value in the time difference indicates that 

Landsat-4 is leading Landsat-5. Landsat-5 leads Landsat-4 

when the value is negative. 

3.2 ORBIT AND GROUNDTRACK MAINTENANCE hANEUVERS 

Following launch and initial groundtrack phasing maneuvers, 

orbit control for the Landsat-5 spacecraft entered the orbit 

maint?nancr? phase. The objectit-e of orbit maintenance is to 

co.~trol the spacecraft altitude within the range that will 

confine the groundtrack to within 10 kilometers cast or west ' 

of the required WRS path. This is accomplished by periodic 

adjustments to the semimajor axis on the order of 100 t? 

300 meters. The rate of semimajor axis decay, and thus the 

period between maintenance maneuvers, depends on the solar 

flux level. The targeted change in semimajor axis will depend on the estimated level of solar activity for several 

months following the maneuver. 

Figure 3-4 illustrates the predicted effect of a typical 

maneuver on groundtrack error depending on the average solar 

flux level encountered (solar flux is given in units of 

watts per square meter per hertz and is for a 

10.7-centimeter wavelength). Following a maneuver, the 

groundtrack begins drifting westward. The objective of the 

maneuver is to force the groundtrack to drift to the western 

boundary (-10 kilometers!, turn around, and drift eastward. 

As the gruundtrack reaches the eastern boundary (+lo kilometers), another maintenance burn is required to reverse the 

drift. If the solar flux level suddenly drops below the 

predicted average soon after the maneuver, the groundt,rack 

may cross the western bour~ary. To correct at tnis point 

would require a retrograde maneuver (180-degree spacecraft 

yaw). Because this is undesirable from a spacecraft operational standpoint, it is necessary to be conservative in 
ORIGINAL PAGE iS 

OF POOR QlJALlTY 
estimating maneuver magnitude. The minimum burn length obtainable with the Landsat-5 propulsion module is 256 milliseconds with two thrusters firing simultaneously. This 

translates into a change in the semimajor axis of approximately 11 meters, which can make a difference of several 

kilometers in the westward groundtrack drift (depending on 

solar activity). The lower the solar flux encountered, the 

greater the difference in westward drift produced by two 

burns differing by one 256-millisecond pulse. Therefore, in 

planning the maneuver, the lowest expected average solar 

flux is used to define an upper limit for the burn time. 

Initial estimates of future solar activity are taken from 

predictions made by the Marshall Space Flight Center. The 

maintenance burns will be inserted in the appendix as they 

occur, beginning with maneuver 9. 

In addition to controlling the groundtrack, the mean local 

time of the descending node must be maintained between 9:30 

and 10:OO a.m. The local time is affected by the nodal regression rate, which is determined by the inclination of the 

orbit. Because the inclination changes slowly in time due 

to the gravitational effects of the Sun and Moon, the nodal 

rate will change causing a change in local ti.. Inclination change maneuvers are required to restore the 

Sun-synchronous node rate. The inclination produced by the 

Delta 3920 launch vehicle was biased in such a way as to 

allow a relatively long time from launch to the first inclination change maneuver. Tt is expected that this maneuver will be required approximately 18 months after launch. 

During the life of the mission, detailed analysis will be 

performed to determine the exact date. Additional maneuvers 

may be required at 8-month intervals thereafter. The change 

in inclination for each burn is expected to be on the order 

of 0.05 degree. The mean local time of the descending node 

at injection was at 9:38 a.m. 
APPENDIX - LANDSAT-5 ORBIT ADJUST MANEUVER DATA 

This appendix contains data and analysis results for all 

maneuvers performed to date. It is intended that new tables 

and figures will be produced as additional maqeuvers are 

performed. For each maneuver, the following two tables and 

two figures are provided: 

Taole of orbit parazeters far the given maneuver 

Table of spacecraft propulsion system parameters 

a Plot of observed groundtrack since the previous 

maneuver 

Plot of observed mean local time of the descending 

node since the previous maneuver 
Table A-1. Orbit Parameters for Maneuver 1 

MANEUVER 1 I ORBIT 83 1 DATE 840307 BURN START TIME (G

PREDICTED POSTBURN OBSERVED POSTBURN 

0 

M 

EPOCH IYYMMDD) 

(HHMMSS) 

AVERAGED ELEMENT!+ I PREBuRN PREDICTED POSTBURN OBSERVED POSTBURN 

(O

EPOCH (YYMMDDI 

(HHMMSS) 

PERIOD (sac) 

PERIGEE ALTITUDEC 

APOGEE ALTITUDE^ 

B COS Z 

I SIN o' 

= SEMIMAJOR AXIS (km) 

e = ECCENTRICITY 

i = INCLINATION Ideg) 

Q = RIGHT ASCENSION OF ASCENDING NODE (degl 

w = ARGUMENT OF PERIGEE (dog) 

M = MEAN ANOMALY (deg) 

TIMES ARE GMT 

MEAN LOCAL TIME OF 

I 093808 

DESCENDING NODE 

IHHMMSS) 

C~QUATO~lAL REFERENCE 
Table A-2. Spacecraft Parameters for Maneuver 1 

MANEUVER ORBIT DATE I BURN START TIME 

IGMT) I 

1 SPACECRAFT PARAMETERS 1 PREBURN I POSTBURN 1 

SPACECRAFT ATTITUDE (degjc 

PITCH 

YAW 

ROLL 

MANEUVER CALIBRATION 

SEMIMAJOR AXIS CHANGE (km) 

PREDICTED 

OBSERVED 

INCLINATION CHANGE ldegl 

PREDICTED 

OBSERVED 

296.94 

17.11 

16.61 

!S. i8 

16.24 

55.61 

55.61 

55.61 

343.36 

510.19 

4284.46 

- - - - - - - - - 

FUEL SYSTEM PRESSURE (PSIA) 297.48 

TANK TEMPERATURES f°C) 

TANK 1 ! 17.11 

THRUST CORRECTION FACTOR 

USED FOR PLANNING 

RECALIBRATED~ 

TANK 2 

TANK 3 

TANK 4a 

HYDRAZlNE REMAINING (POUNDS) 

TANK 1 

TANK 2 

TANK 3 

TANK 4 

TOTAL FUEL 

TOTAL SPACECRAFT WEIGHT 

"ANK 4 IS THE AUXILLIARY TANK KIT (ATK) 

%URN TIME INPUT TO GENERAL MANEUVER PROGRAM (GMAN) =TOTAL DURATION - NUMBER OF THRUSTERS 

C~P~C~~~~~T ATTITUDE AS INPUT TO GMAN FOR MANEUVER MODELING 

*RECALIBRATED THRUST CORRECTION FACTOR = IOBSERVED - PREDICTED) x FACTOR USED FOR PLANNING 

. 16.6: 

18.18 

16.24 

55.67 

55.67 

55.67 

343.50 

510.51 

4284.78 

THRUSTERS 

ORBIT ADJUST THRUSTERS USED 

TOTAL ORBIT ADJUST THRUSTER DURATION (secjb 

TOTAL AlTlTUDE THRUSTER DURATION lsec) 

81. Dl 

9.216 

102.OBO 

CRIGlNAL PAGE I9 

OF POOR QUALITY 
Table A-3. Orbit Parameters for Maneuver 2 

a 

e 

i 

Q 

W 

M 

EPOCH 

MANEUVER 

(YYMMDD) 

IHHMMSS) 

2 

OSCULATING ELEMENTS= 

AVERAGED ELEMENTS~ I PREBURN 

- 

I 

5 

- 

0 

G 

EPOCH IYYMMOD) 

(HHMMSSI 

PERIOD (sac) 

PERIGEE ALTITUDE^ 

APOGEE ALTITUDE^ 

B COS Tj 

I SIN o' 

GROUNDTRACK ERROR (kmld 

ORBIT 119 

PREBURN 

MEAN LOCAL TIME OF 

I 093817 

DESCENDING NODE 

(HHMMSSJ 

- -- 

PREDICTED POSTBURN 

- 

7066.4960 

0.0002649 

90.2548689 

132.31769 

175.26228 

285.73644 

840309 

212651.608 

5911.76 

686.484 1 

690.2279 

- 0.0002640 

0.0000219 

DATE 840309 

PREDICTED POSTBURN 

7057.9823 

1 0.0015186 

i 98.2589719 

I 

I 132.31973 

I 277.25790 

183.72113 

I 840309 

212651.608 

OBSERVED POSTBURN 

7088.4964 

0.0002658 

98.2549813 

132.31773 

175.C2668 

285.37219 

840309 

212651.608 

M11.75 

686.4681 

690.2247 

- 0.0002650 

0.0000203 

BURN START TIME IGM

OBSERVED POSTBURN 

1 = SEMIMAJOR AXIS (km) 

I e = ECCENTRICITY 

i = INCLINATION (deg) 

(OB

I I Q = RIGHT ASCENSION OF ASCENDING NODE (deg) 

w = ARGUMENT OF PERIGEE (deg) 

M = MEAN ANOMALY (deg) 

TIMES ARE GMT 
Table A-4. Spacecraft Parameters for Maneuver 2 

I MANEUVER I ORBIT I DATE I BURN START TIME 

IGMT) I 

1 SPACECRAFT PARAMETERS ( PREBURN I POSTBURN I -- -- - - - 

FUEL SYSTEM PRESSURE (PSIA) 

TANK TEMPERATURES (OC) r TANK I 

SPACECRAFT ATTITUDE (deglc 

PITCH 

YAW 

ROLL 

TANK 2 

TANK 3 

TANK 4a 

HYDRAZINE REMAINING (POUNDS) 

TANK 1 

TANK 2 

TANK 3 

TANK 4 

TOTAL FUEL 

TOTAL SPACECRAFT WEIGHT 

MANEUVER CALIBRATION 

SEMIMAJOR AXIS CHANGE (km) 

P3EDICTED 

OBSERVED 

INL-INATION CHANGE Ideg) 

PREDICTED 

OBSERVED 

17.11 

18.71 

16.23 

55.61 

55.61 

56.61 

343.36 

510.19 

4284.46 

aTAN~ 4 IS THE AUXILLIARY TANK KIT fATK) 

%URN TIME INPUT TO GENERAL MANEUVER PROGRAM (GMAN) =TOTAL DURATION - NUMBER OF THRUSTERS 

CSPACEC~AFT ATTITUDE AS INPUT TO GMAN FOR MANEUVER MODELING 

 REC CALIBRATED THRUST CORRECTION FACTOR =(OBSERVED - PREDICTED) x FACTOR USED FOR PLANNING 

17.11 

18.71 

16.23 

56.57 

55.V 

56.57 

343.26 

509.97 

4284.24 

THRUSTERS 

ORBIT ADJUST THRUSTERS USED 

TOTAL ORBIT ADJUST THRUSTER DURATION (seclb 

TOTAL ATTITUDE THRUSTER DURATION Isec) 

THRUST CORRECTION FACTOR 

USED FOR PLANNING 

RECALIBRATED~ 

81, Dl 

9.216 

7.840 

1 .MX)O 

0.9~~38 


,.*- 

Table A-5. Orbit Parameters for Maneuver 3 

MANEUVER 3 

OSCULATING ELEMENTS~ 

a 

e 

i 

Q 

w 

M 

EPOCH IYYMMDD) 

(HHMMSS) 

- 

a 

- 

e 

- 

i 

B 

0' 

- 

M 

EPOCH (YYMMDD) 

(HHMMSS) 

PERIOD lsacl 

PERIGEE ALTITUDE' 

APOGEE ALTITUDE^ 

I COS T; 

B SIN Ei -- -- 

GROUNDTRACK ERROR (kmId 

MEAN LOCAL TlME OF 

DESCFNDING NODE 

IHHMMSS) 

ORBIT 192 

PREBURN 

7058.5342 

0.0013113 

98.2595169 

137.30470 

271.67404 

193.41979 

840314 

212918.888 

PREBURN 

7066.4749 

0.0005825 

98.254866a 

137.30203 

167.29515 

297.82554 

840314 

212918.888 

5911.74 

684.2187 

692.451 1 

- 0.0006682 

0.0001281 

DATE 840314 I BURN START TIME 

-- -- 

PREDICTED POSTBURN I OBSERVED POSTBURN I -- 

7059.9786 

0.0011094 

98.2594774 

137.30485 

269.91821 

195.17410 

840314 

212918.888 

PREDICTED POSTBURN 

7067.9243 

0.L006816 

98.2Yrn79 

137.30218 

151.55642 

313.56275 

840314 

212918.888 

5913.56 

684.8961 

694.6725 

- 0.0006081 

0.0003294 

-- -- 

OBSERVED I 

aa = SEMIMAJOR AXIS (kml 

e = ECCENTRICITY 

i = INCLINATION (d-I 

Q = RIGHT kSCENSlON OF ASCENDING NODE (degl 

w = ARGUMENT OF PERIGEE (-I 

M = MEAN ANOMALY (deg) 

TIMES ARE GMT 
Table A-6. Spacecraft Parameters for Maneuver 3 

I 

MANEUVER I ORBIT 

SPACECRAFT PARAMETERS 

FUEL SYSTEM PRESSURE (PS!Al 

TANK TEMPERATURES (OC) 

T4NK 1 

TANK 2 

TANK 3 

TANK 4a 

HYDRAZINE REMAINING (POUNDS) 

TANK 1 

TANK 2 

TANK 3 

TANK 4 

TOTAL F .EL 

TOTAL SPACECRAFT WEIGHT 

DATE 

PREBURN 

THRUSTERS 

ORBIT ADJUST THRUSTERS USED 

TOTAL ORBIT AD~UST THRUSTER DURATION (sec)b 

TOTAL ATTITUDE TIIRUSTER DURATION (secl 

SPACECRAFT ATTITUDE (deglc 

PITCH 

YAW 

ROLL 

MANEUVER CALIBRATION 

SEMIMAJOR AXIS CHANGE (kml 

PREDICTED 

OBSERVED 

INCLlI:ATION CHANGE (degl 

PREDICTED 

OBSESVED 

THRUST CORRECTION FACTOR 

USED FOR PLANNING 

RECALIBRATED~ 

BURN ST.4RT TIME 

(GMT; 

-- - 

POSTBURN 

Al, 01, C1, Dl 

63.552 

17.080 

N/ A 

NIA 

a~~~~ 4 IS THE AUXILLIARY TANK KIT (ATK) 

 BURN TIME INP~JT TO GENERAL MANEUVER PROGRAM (GMANI =TOTAL DURATION + NUMBER OF THRUSTERS 

C~~~~~rS~FT ATTITUDE AS INPUT TO GMAN FOR MANEUVEBJvlOOEUNG-, . -- 

P RE CALIBRATED THRUST CORRECTION FACTOR ~(OBSERVED + PREDICTED) x FACTOR USED FOR PLANNING 

Q":. '".*RL ;"I:.'.. ' ig a:L.l:d , k 

OF POOR QUALITY 
Table A-7. Orbit Parameters for Maneuver 4 

MANEUVER 4 

OSCULATING ELEMENTS~ 

0 

M 

EPOCH (YYMMDDI 

IHHMMSSI 

- 

a 

- 

e 

- 

i 

Si 

- 

W 

ii 

EPOCH (YYMMDDI 

IHHMMSS) 

PERIOD Isecl 

PERIGEE ALTITUDE' 

APOGEE ALTITUDE^ 

z COS 0 

T SIN 'i 

I MEAN LOCAL TIME OF 

DESCENDING NODE 

(HHMMSS) 

- 

ORBIT 404 

PREBURN 

PREBURN 

DATE 840329 

PREDICTED POSTBURN 

PREDICTED POSTBURN 

= SEMIMAJOR AXIS (kml 

e = ECCENTRICITY 

i = INCLINATION Ideg) 

Q = RIGHT ASCENSION OF ASCENDING NODE (dngI 

o = ARGUMENT OF PERIGEE (deg) 

M = MEAN ANOMALY (deg) 

TIMES ARE GMT 

BURN START 

OBSERVED POSTBURN 

7072.5850 

0.0014319 

98.2497870 

151.81153 

152.10850 

174.11100 

840329 

1 10946.977 

OBSERVED 

7069.0616 

0.0012524 

98.2520559 

151 .SO653 

134.37762 

191.891 68 

840329 

110946.977 

5815.01 

682.0083 

699.7949 

- 0.0008759 

00008951 

'EQUATORIAL REFERENCE 

TIME (GMT

IOBSE

- 

- 

- 
Table A-8. Spacecraft Parsme* rs for Maneuver 4 

I MANEUVER I ORBIT I DATE I BURN START TIME 

I IGMT) I 

I 

- 

SPACECRAFT PARAMETERS I PREBURN I POSTBURN I - - -- - - 

FUEL SYSTEM PRESSURE (PSIA) 

TANK TEMPERATURES (OC) 

TANK 1 

TANK 2 

TANK 3 1 1B.W 1 18.00 

TANK 4a I 14.26 1 14.26 

HYDRAZINE REMAINING (POUNDS) 

TANK 1 I S.30 

TANK 2 55.30 

TANK 3 55.30 

TANK 4 I 342.57 

TOTAL FUEL I 508.47 1 507.19 

ORBIT ADJUST THRUSTERS USED I Al, 81. C1. Dl 

TOTAL SPACECRAFT WEIGHT 

TOTAL ORBIT ADJUST THRUSTER DURATION (secP 

TOTAL ATTITUDE THRUSTER DURATION (sec) 

4282.74 I 4281.46 

SPACECRAFT AlTlTUDE (deglc 

PITCH 

YAW 

ROLL 

THRUSTERS 

MANEUVER CALIBRATION 

SEMIMAJOR AXIS CHANGE (km) 

PREDICTED 

OBSERVED 

INCLINATION CHANGE (deg) 

PREDICTED 

OBSERVED 

NIA 

Nl A 

a~~~~ 4 IS THE AUXlLLlARY TANK KIT (ATKI 

THRUST CORRECTION FACTOR 

USED FOR PLANNING 

RECALIBRATED~ 

%URN TIME INPUT TO GENERAL MANEUVER PROGRAM (GMANI =TOTAL DURP.TION-NUMBER OF THRUSTERS 

CSP~C~CRAFT AlTlTUDE AS INPUT TO GMAN FOR MANEUVER MODELING 

 REC CALIBRATED THRUST CORRECTION FACTOR =(OBSERVED + PREDICTED) x FACTOR USED FOR PLANNING 

1.0000 

0.9988 

HOURS IN NODE DESCENDING OF TIME LCCAL E4Y 

' z i 0 

m l t" n&r 1 

1 0 

I 

' 4 0 

I., - 3 I n, 

c.n ~ b 

i ' -.
Table A-9. Orbit Parameters for Maneuver 5 

I MANEUVER 5 I ORBIT 405 

( OSCULATING ELEMENTS~ I PREBURN 

a 

e 

I 

Q 

W 

M 

EPOCH IYYMMDD) 

(HHMMSSI 

I AVERAGED ELEMENTS~ I PREBURN 

- 

a 

- 

e 

- 

i 

Ti 

- 

w 

fi 

EPOCH IVYMMDD) 

(HHMMSS) 

PERIOD Isec) 

PERIGEE ALTITUDE' 

APOGEE ALTITUDE' 

Z COS i2 

Z SIN C 

GROUNDTRACK ERROR lkrnld 

DATE 840329 BURN START TIME (GM

7069.0767 

0.0012692 

SB.2511642 

151 .a7381 

134.78209 

187.55100 

840329 

124724.336 

5915.00 

681.9646 

699.9088 

- 0.0008940 

0. oo09009 

- 66.8 

I MEAN LOCAL TIME OF 

DESCENDING NODE 

(HHMMSSI 

PREDICTED POSTBURN I OBSERVED POSTBURN I 

093905 

PREDICTED POSTBURN I OBSERVED POSTBURN 1 IOB

= SEMIMAJOR AXIS (km) 

e = ECCENTRICITY 

I = INCLINPTION Idegl 

Q = RIGHT ASCENSION OF ASCENDING NODE IdegI 

w = ARGUMENT OF PERIGEE Idegl 

M = MEAN ANOMALY (deg) 

TIMES ARE GMT 

'EQUATORIAL REFERENCE 
Table A-10. Spacecraft Parameters for Maneuver 5 

MANEUVER 

FUEL SYSTEM PRESSURE (PSIAI 

TANK TEMPERATURES (OCI 

TANK 1 

TANK 2 

TANK 3 

TANK 4a 

HYDRAZINE REMAINING (POUNDS) 

TANK 1 

TANK 2 

TANK 3 

TANK 4 

TOTAL FUEL 

TOTAL SPACECRAFT WEIGHT 

THRUSTERS 

ORBIT ADJUST THRUSTERS USED 

TOTAL ORBIT ADJUST THRUSTER DURATION (see+ 

TOTAL AlTITUDE THRUSTER DURATION (oecl 

ORBIT I 

SPACECRAFT ATTITUDE (deglc 

PITCH 

YAW 

ROLL 

840329 

PREBURN 

5 

MANEUVER CALIBRATION 

SEMIMAJOR AXIS CHANGE (km) 

PREDICTED 

OBSERVED 

INCLINATION CHANGE ldegl 

PREDICTED 

OBSERVED 

DATE 

124654 

POSTBURN 

405 

THRUST CORRECTION FACTOR 

USED FOR PLANNING 

RECALIBRATED~ 

BURN START TIME 

(GMTI 

SPACECRAFT PARAMETERS 

"ANK 4 IS THE AUXlLLlARY TANK KIT (ATKI 

%URN TIME INPUT TO GENERAL MANEUVER PROG~AM (GMANI =TOTAL DURATION - NUMBER OF THRUSTERS 

CSP~CEC~~FT ATTITUDE AS INPUT TO GMAN FOR MANEUVER MODELING 

 REC CALIBRATED THRUST CORRECTION FACTOR =(OBSERVED - PREDICTED) x FACTOR USED FOR PLANNING 


Table A-11. Orbit Parameters for Maneuver 6 

1 MANEUVER 6 1 ORBIT 425 1 DATE 840330 

I OSCULATING ELEMENTS~ I PREBURN 1 PREDICTED POSTBURN 

a 

e 

i 

Q 

W 

M 

EPOCH IYYMMDD) 

I AVERAGED ELEMENTS~ I PREBURN I PREDICTED POSTBURN 

BURN START 7 

OBSERVEDPOSTBURN 

e' 

- 

i 

5 

- 

w 

fi 

EPOCH (YYMMDD) 

IHHMMSSI 

PERIOD IsecJ 

PERIGEE ALTITUDE^ 

APOGEE ALTITUDE^ 

5 COS Ci 

T SIN i3 

[ GROUNDTRACK ERROR (krnld I 18.4 'a = SEMIMAJOR AXIS (km) 

214843.928 

OBSERVED 

0.0010242 

98.2523636 

153.24053 

131.11557 

214.19665 

840330 

214843.928 

5917.88 

685.9867 

700.4717 

- 0.0006735 

0.000n16 

I MEAN LOCAL TIME OF 

I 093908 

DESCENDING NODE 

(HHMMSS) 

(OBS

0.0007998 

98.2526012 

153.24046 

121.50841 

223.80147 

840330 

214843.928 

5920.28 

689.4842 

700.7986 

- 0.0004180 

0.m19 

I e = ECCENTRICITY 

i = INCLINATION Ideg) I I 9 = RIGHT ASCENSION OF ASCENDING NODE IdegJ 

w = ARGUMENT OF PERIGEE Idegl 

M = MEAN ANOMALY Ideg) 

TIMES ARE GMT 

CEQUATORl~L REFERENCE 
Table A-12. Spacecraft Parameters for Maneuver 6 

b6 - 1 

MANEUVER 

6 

bB~~~ TIME INPUT TO GENERAL MANEUVER PROGRAM (GMAN) =TOTAL DURATION + NUMBER OF THRUSTERS 

CS~~C~~~~~T AlTlTUDE AS INPUT TO GMAN FOR MANEUVER MODELING 

~RECALIERATED THRUST CORRECTION FACTOR =(OBSERVED + PREDICTED) x FACTOR USED FOR PLANNING 

BURN START TIME 

(GMT) 

214803 

ORBIT 

425 

POSTBURN 

281.06 

17.11 

16.40 

18.06 

13.95 

54.26 

54.26 

54.26 

339.92 

502.70 

4276.97 

DATE 

840330 

SPACECRAFT PARAMETERS 

FUEL SYSTEM PRESSURE (PSIA) 

TANK TEMPERATURES (OC) 

TANK 1 

TANK 2 

TANK 3 

TANK 4a 

HYDRAZlNE REMAINING (POUNDS) 

TANK 1 

TANK 2 

TANK 3 

TANK 4 

TOTAL FUEL 

TOTAL SPACECRAFT WEIGHT 

PREBURN 

284.24 

17.11 

16.40 

18.06 

13.95 

54.62 

54.62 

54.62 

340.82 

504.68 

4278.95 

THRUSTERS 

ORBIT ADJUST THRUSTERS USED 

TOTAL ORBIT ADJUST THRUSTER DURATION (seclb 

TOTAL ATTITUDE THRUSTER DURATION (set) 

SPACECRAFT ATTITUDE ldeglC 

PITCH 

YAW 

ROLL 

MANEUVER CALIBRATION 

SEMIMAJOR AXIS CHANGE (krn) 

PREDICTED 

OBSERVED 

INCLINATION CHANGE (degl 

PREDICTED 

OBSERVED 

THRUST CORRECTION FACTOR 

USED FOR PLANNING 

RECALIBRATED~ 

Al, C1 

81.856 

87.360 

0.0 

0.0 

0.0 

1.9123 

1 3626 

N/ A 

N/A 

1.0000 

0.9739 - 

a~~~~ 4 IS THE AUXlLLlARY TANK KIT (ATK) 

ORIGINAL PAGE is 

OF POOR QUALITY 

IS; 7. 
Table A-13. Orbit Parameters for Maneuver 7 

C~9~~~O~~~~ REFERENCE 

t' 

I\) 

cn 

w .= ARGUMENT OF PERIGEE Ldeg) 

M = MEAN ANOMALY Ldeg) 

TIMES ARE GMT 

MANEUVER 7 

OSCULATING ELEMENTS~ 

a 

e 

i 

Q 

w 

M 

EPOCH (YYMMDD) 

IHHMMSS) - 

AVERAGED ELEM~NTS~ 

- 

a 

- 

e 

- 

I 

G 

- 

w 

- 

M 

EPOCH IYYMMDDI 

(HHMMSF) 

PERIOD lsecl 

PERIGEE ALTITUDE= 

APOGEE ALTITUDE^ 

B COS ij 

i S!N G 

GROUNDTRACK ERROR (kmld 

MEAN I CCAL TIME OF 

DESCcNDlNG NODE 

IHHivlMSS) 

ORBIT 469 

PFtEBURN 

7075.6386 

0.0013429 

98.2492276 

156.17316 

136.31083 

259.32621 

840402 

205102.432 

PREBURN 

7073.2325 

0.0009096 

98.2505715 

156.17840 

122.67551 

274.9CB97 

840402 

205102.432 

5920.22 

688.6587 

701.5263 

-- 0.00491 1 

O.OW7656 

- 38.0 

093913 

. . - -- - 

DATE 840402 

PREDICTED POSTBURN 

7078.3570 

0.0012923 

98.2495193 

156.17337 

121 56387 

275.96968 

840402 

-. 205102.432 

PREDICTED POS TBURN 

7075.9524 

0.000986ti 

98.2- 

156.17S1 

99.73043 

297.85057 

840402 

205102.432 

5923.63 

690.8320 

704.79'28 

aa = SEMIMAJOR AXIS (km) 

e = ECCENTRICITY 

i = INCL'NATION (deg) 

Q = RIGHT ASCENSION OF ASCENDING NODE (deg) 

- 0.0001667 -0.0001KQ 

O.Ml09723 I 0.0MH716 

BURN START 

OBSERVEDPOSTBURK 

7078.2720 

0.0012989 

98.2499165 

156.17369 

122.18392 

775.45000 

840402 

205102.432 

OBSERVED 

175.8673 

0.0009882 

98.2512603 

156.17894 

100.50609 

297 07528 

840402 

205102.432 

5823.53 

690.7349 

704.7197 

TIME (GMT

LOBS
Table A-14. spacecraft Parameters for Maneuver 7 

I MANEUVER ( ORBIT DATE I BURN START TIME 

(GMT) I 

7 

SPACECRAFr ATTITUDE (deglc 

PITCH 

YAW 

ROLL 

TANK 2 

TANK 3 

TANK 4a 

HYDRAZINE REMAINING (POUNPS) 

TANK 1 

TANK 2 

TANK 3 

TANK 4 

TOTAL FUEL 

TOTAL SPACECRAFT WEIGHT 

MANEUVER CALIBRATION 

SEMIMAJOR AXIS CHANGE (km) 

PREDICTED 

OBSERVED 

INCLINATION CHANGE Ideg) 

PREDICTED 

OBSERVED 

469 

16.39 

17.87 

13.94 

54.26 

54.26 

54.26 

339.92 

502.m 

4276.97 

Al, C1 

116.864 

111.160 

SPACECRAFT PARAMETERS 

FUEL SYSTEM PRESSURE (PSIA) 

TANK TEMPERATURES (OC) 

TANK 1 

THRUSTERS 

ORBIT ADJUST THRUSTERS USED 

TOTAL ORBIT ADJUST THRUSTER DURATION rsec~b 

TOTAL AlTlTUDE THRUSTER DURATION (sec) 

THRUST CORRECTION FACTOR 

USED FOR PLANNING 

840402 

a~~~~ 4 IS THE AUXlLLlARY TANK KIT IATK) 

%URN TIME INPUT TO GENERAL MANEUVER PROGRAM (GMAN) =TOTAL DURATION - NUMBER OF THRUSTERS 

C~?~CECR~FT ATTITUDE AS INPUT TO GMAN FOR MANEUVER MODELING 

~PECALIBRATED THRUST CORRECTION FACTOR = (OBSERVED - PREDICTED) x FACTOR USED FOR PLANNING 

206004 

PREBURN 

280.89 

16.94 

POSTBURN 

276.54 

16.94 
ORIGINAL PAGE IS 

OF POOR QUALITY 

Table A-15. Orbit Parameters for Maneuver 8 

MANEUVER 8 

OSCULATING ELE~ENTS~ 

a 

e 

i 

Q 

0 

M 

EPOCH (YYMMDDI 

IHHMMSSI 

EPOCH (YYMMDDI 

IHHMMSSI 

PERIOD Isecl 

PERIGEE ALTITUDE' 

APOGEE ALTITUDE' 

6 COS & 

V SIN B 

GROUNDTRACK ERROR (kmld 

MEAN LOCAL TlME OF 

DESCENDING NODE 

(HHMMSSI 

ORBIT 498 DATE 840404 I BURN START TIME 

PREBURN PREDICTED POSTBURN OBSERVED POSTBURN 

PREBURN 

I PREDICTED POSTBURN 1 OBSERVED WSTBURN I 

+I aa = SEMIMAJOR AXIS (kml 

e = CCCENTRICITY 

i = INCLINATION (degl 

Q = RIGHT ASCENSION OF ASCENDING NODE (deg) 

w = ARGUMENT OF PERIGEE (deg) 

M MEAN ANOMALY IdegI 

TIMES ARE GMT 

b~~~AVERAGED OVER ONE ORBIT 

C~QUATORl~~ REFERENCE 
Table A-16. Spacecraft Parameters for Maneuver 8 

( MANEUVER I ORBIT I DATE I BURN START IGMT) I 

I SPACECRAFT PARAMETERS I PREBURN I WSTBURN I 

FUEL SYSTEM PRESSURE (PSIAI 

TANK TEMPERATURES i°C) 

TANK 1 

TANK 2 

TANK 3 

TANK 4a 

F YDRAZINE REMAINING i POUNDS) 

TANK 1 

TANK 2 

TANK 3 

TANK 4 

TOTAL FUEL 

TOTAL SPACECRAFT WEIGHT 

THRUSTERS 

ORBIT ADJUST THRUSTERS USED 

TOTAL ORBIT ADJUST THR~~STER DURATION isec~b 

TOTAL AnTUDE THRUSTER DURATION isec) 

SPACECRAFT AlTlTUDE ideglc 

PITCH 

YAW 

ROLL 

MANEUVER CALIBRATION 

SEMIMAJOR AXIS CHANGE (km) 

PREDICTED 

OBSERVED 

INCLINATION CHANGE (deg) 

PREDICTED 

OBSERVED 

Al, C1 

86.336 

83.440 

N/A 

NIA 

aT~~~ 4 IS THE AUXlLLlARY TANK KIT IATK; 

THRUST CORRECTION FACTOR 

USED FOR PLANNING 

RECALIBRATED~ 

b~~~~ TIME INPUT TO GENERAL MANEUVER PROGRAM iGMANl = rOTAL GURATION -NUMSER OF THRUSTERS 

C~~~C~~~~FT ATTITUDE AS INPUT TO GMAN FOR MANEUVER MODELING 

 REC CALIBRATED THRUST CORRECTION FACTOR =(OBSERVED - PREDICTED) x FACTOR USED FOR PLANNING 

0.9700 

0.9624 
(J~~GINAL PfiC'Z !g 

OF POOR QUALln 

REFERENCES 

1. Computer Sciences Corporation, CSC/TM-82/6232, Landsat-4 

Orbit Adjust Maneuver Report, R. J. McIntosh, 

December 1982 

2. General Electric Space Division, Landsat-D Orbit Adjust 

Criteria, P. Jasper, February 1982 

3. Computer Sciences Corporation, CSC/TM-79/6089, Description and Application of the Frozen Orbit Concept, 

R. W. ~erd3 1979 

4. -- , CSC/TM-82/6019, Landsat-D Propulsion System Modeling 

and Maneuver Targeting, R. J. McIntosh, April 1982