Results are presented of a study to show the effect of selecting a constant inertial attitude during the fuel dissipation phase of a return-to-launch-site abort. Results are also presented which show that the selection of the constant inertial attitude will affect the arrival point on the range-velocity target line. An alternate selection of the inertial attitude will provide control over the trajectory shape

Bown, R. L.

Winans, L. C.

English

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Produced by the NASA Center for Aerospace Information (CASI) 
https://ntrs.nasa.gov/search.jsp?R=19760026148 2020-03-22T13:49:40+00:00Z
I.
MCDONNELL DOUGLAS TECHNICAL SERVICES CO.
HOUSTON ASTRONAUTICS DIVISION
SPACE SHUTTLE ENGINEERING AND OPERATIONS SUPPORT
DESIGN NOTE NO. 1.4-4-7
RETURN-TO-LAUNCH-SITE THREE DEGREE OF FREEDOM ANALYSIS,
CONSTANT INERTIAL ATTITUDE DURING THE FUEL DISSIPATION PHASE
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PREPARED BY: L ^ li^,ttt rcaa
L. C. Winans
Task Manager
488-5660, Ext. 2.43
APPROVED BY: 1!,.Cj
^;. W. F. Hayes
Project F?anager
Mission Planning,
Fission Analysis and
Softv.,are Formulation
4£s-66CO, Ext. 266
PREPARED BY: 2a"'/
R. L. Bown
Senior Engineer
483-5660, Ext. 243
APPROVED BY:/
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FF'3 t •lork Package t:.Enage ►
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This Design Note is Submitted to NASA Under Task Order
No. D0303, Task Assignment 1.4-4-D in Fulfillment of
Contract NAS 9-13970.
EJ	 ,
1 	 SUIWRY
This document presents the results of a study to show the effect
of selecting a constant inertial attitude during the fuel dissipation
i
phase of a Return-to-Launch-Site (RTLS) abort. Results are presented
which show that the selection of the constant inertial attitude will
affect the arrival point on the Range-Velocity (R-V) target line.
An alternate selection of the inertial attitude will provide control
over the trajectory shape.
At
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DN No.: 1.4-4-1
Page: 2
2.0 INTRODUCTION
Preliminary RTLS guidance and targeting software for the Space
Shuttle is documented in Reference (A). This note documents the
first of a series of performance verification studies planned to
verify the adequacy of that software.
After a main engine shutdown and crew selection of the RTLS mode,
a fuel dissipation phase subsequent to solid rocket booster (SRB)
staging is required. The duration of this phase is greatest for
early aborts and decreases to zero near the mode boundary. A
constant inertial attitude during this phase will affect the
trajectcry and the point of arrival on the R-V line. The purpose
of this note is to parametrically examine the amount of R-V line
and trajectory control available.
One goal of the trajectory control is to make the flyback trajectories
neighboring. This may be important from the monitoring and/or
reversion to manual back-up flight.
DN No.: 1.4-4-7
Page: 3
3.0 DISCUSSION
This study was conducted to determine the changes in the arrival
points on the main engine cutoff (MECO) R-V line and in the fly-
back trajectories caused by changes in the constant inertial
attitude. After a space shuttle main engine (SSME) failure and
an RTLS command, the space shuttle is rotated to a predetermined
attitude and continues downrange to dissipate fuel at this constant
inertial attitude. The abort times used in this study were 140,
180, and 220 seconds frorr, launch. The range of thrus t, directions
or body attitudes (which are coincident in the 3 degree of freedom
simulation) was 40 to 60 degrees measured from a plane normal to
the local geodetic vertical at time of launch (Figure 1). This
reference plane is fixed for the entire RTLS simulation. The
rotation to the desired attitude is done in the pitch plane or
about the body axis through the orbiter wings.
This study used a three degree of freedom simulation contained on
a modified Space Vehicle Dynamic Simulation (SVDS) 2.3.11 milestone
file (Reference (B)) for a mission 3A RTLS abort launched from the
Western Test Range. The modifications to SVDS were:
a) Addition of the turnaround time prediction logic (Reference (C)).
b) Addition of the thrust termination logic (reference (C)).
The inputs to PEG were biased tc the Main Fngine Cutoff minus ten
seconds (MECO-10) target conditions, total weight of 310,000 pounds,
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DN No.: 1.4-4-7
Page: 5
1
230,000 feet altitude, and a 4 degree earth relative flight path
angle. The biased desired flight path angle at the MECO-10 R-V
target line results in an angle near zero at external tank
separation. The Rockwell International (RI) R-V target line for
MECO-10 was used:
R - .069VR - 110.1
For thrust termination the target was the RI 11ECO R-V line:
R = .068VR - 171.5
All ranges are from the landing site at the Western Test Range in
nautical miles.
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DN No.: 1.4-4-7
Page: 6
4.0 RESULTS
Typical thrust direction histories for aborts occurring at 140 and
i
220 seconds are shown in Figures (2) and (3). The variation in
termination conditions at external tank separation is shown below:
altitude
	
229,126 to 230,763 feet
flight path angle 	 .15 to .75 degrees
dynamic pressure	 3.94 to 4.63 p.s.f.
More precise targeting and prediction of powered pitch down effects
on MECO conditions could reduce the variation but would not be
meaningful at this stage in the performance testing.
An examination of the i:umerical results presented in Tables I to III
and Figures (2) and (3) provides the following general observations.
The time of abort has a definite effect upon the arrival point on
the R-V line (i.e. the earlier the abort the higher the terminal
velocity). A thrust direction of 60 degrees tends to loft the
trajectory such that the space shuttle achieves altitudes in excess
of 400,000 feet with an associated lower velocity.
The turnaround time shown in the abort time tables are the times at
the end of the downrange fuel dissipation phase. The choice of down-
range gimbal angle causes a difference of up to 43.5 seconds in the
turnaround time with the nose low case turning earlier. Essentially,
at turnaround tire the lofted trajectory has a higher altitude, lower
velocity, greater range, and less fuel remaining than the 40 degree
DN No.: 1.4-4-7
Page: 7
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Page: 9
TABLE I
Conditions at Selected Trajectory Study Points
Constant Inertial Thrust Direction
140 Second Abort
CONDITION THRUST DIRECTIONS
600	500	 400
Turnaround
T;^;,e -
	
sec 321.7 301.1 278.2R
,.,- ative Velocity - fps 7453.3 7744.2 7757.0
Range - n.m. 229.24 213.40 190.76
Altitude - feet 440305 372271 325082
MECO - 10
Weight - lbs 309671 309354 309839
Altitude - feet 224756 224442 224262
Flight Path Angle - deg 3.939 3.973 4.004
MECO
Range - n.m. 293.78 308.15 315.63
Relative Velocity - fps 6839.4 7052.8 7161.5
Time - sec 646.7 F41.3 639.1
Turnaround
Time - sec 292.6
Rela',.ive Velocity - fps 7847.6
Range - n.m. 212.70
MECO - 10
Weight - lbs 309755
Altitude - feet 224530
Flight Path Angle - deg 3.935
MECO
Range - n.m. 293.84
Relative Velocity - fns 68G1.0
Time - sec 626.6
267.0
8088.6
186.03
309864
224481
3.994
305.96
7027.7
(120.1
'179.3
8032.9
199.52
309853
224.564
3.975
301.50
6954.6
622.7
f
DN No.: 1.4 -4-7
Page: 10
TABLE II
Conditions at Selected Trajectory Study Points
Constant Inertial Thrust Direction
180 Second Abort
I I
CONDITION
	
THRUST DIRECTIONS
60°
	
50°
	
40°
TABLE III	
DN No.: 1.4-4-7
Page: 11
Conditions at Selected Trajectory Study Points
Constant Inertial Thrust Direction
220 Second Abort
1
CONDITION THRUST DIRECTION
60° 500 40°
Turnaround
Time - sec 260.8 255.4 250.9
Relative Velocity - fps 8304.4 8360.8 8379.7
Range - n.m. 182 .97 176.14 170.50
Altitude - feet 353264 345255 338C85
MECO - 10
Weight - lhs 309746 309810 309877
Altitude - feet 224578 224613 224691
Flight Path Angle - deg 3.963 3.977 4.000
M000
Range - n.m. 291,71 293.37 295.30
Relative Velocity - fps 6809.4 h°35,7 F864.1
Tire - sec 605.1 603.1 600.8
of TO
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DV No.: 1.4-4-7
Paye: 12
trajectory. The lofted trajectory overtakes the lower trajectory
and terminates lower on the R-V line.
Figure (4) is a carpet plot which shows a sumr^ilary of the effect of
inertial attitude during the fuel dissipation phase and abort time
on range to launch site at MECO. Since all cases terminated at
the R-V line, the ii lative velocity associated with the range is
available from th, MECO R-V expression presented above. The velocity
spread at the R-V line, for the inertial attitudes considered; range
from 322 ft/sec for the earliest aborts to 55 ft/sec for a 220 sec.
The spread would continue to decrease until 235 . 240 second
abort when control over the point of arrival on the R-V line mould
no longer be available. This abort time is, of course, not the
latest available but represents the latest for flyback at 100
percent throt t le. The RTLS/Abort Once Around (AOA) boundary is at
approximately 255 seconds for flyback at 109 percent throttle.
One of the reasons for control over the velocity at the f,-V line is
to achieve more ber.irn entry conditions. That is, a higher velocity
results in h i gher dynamic pressures which tend to alleviate cntry
load relief problems in that the uncle of attack need riot tic , rcduced
so much for 9 limiting and thus the excursion into th(- less c'E sir< 1,1e
;)UCll3IL1TY OF TR,r
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320
DN No.: 1.4-4-7
FIGURE 4	 Page: 13
DECO R-V Arrival Points
Abort Time and Fuel Dissinatinn	 40°
2 0
	
285	 290	 295	 300	 305	 310	 315
Range in nautical miles
.',It0DUCIBILITY OF
ANIAL PAGE IS PC("",
DN No.: 1.4-4-7
Page: 14
stability area is limited. However, the velocity spread available
only results in a maximum of 9% change in dynamic pressures at
separation. A 5000 ft decrease in altitude on the other hand
results in a 25% increase. It would appear, subject to further
investigation, that altitude variations for dynamic sressure control
would be a more profitable approach. Another use for the velocity
control would be to arrive at a "point" on the R-V line for all
abort times thus making the entry basically the same for each
mission and reducing crew training requirements for the entry phase.
Referring, i^•_ain to figure (4) it can be seen that there is no one
range (and therefore velocity) that.is compon to all abort times
although the range of entry conditions could certainly be narrowed.
Range control is more significant tirith 21.9 Nfi available at the
earliest aborts, 4.6 miles at 220 sec. abort and, again, falling
to zero at approximately 235-240 second aborts. Range control is
not felt to he significant in the sense that each R-V point
represents a point of equal opportunity for successful return. It
is thought that a better approach world be to translate the target
R-V line towards the launch site when the abort time is earlier
than 235-240 seconds and, by definition excess fuel reeiains. This
results in a "pad" or arrival nearer the center of the Terminal
Area Energy Management (TAL III) footprint. Targeting to the center
of the footprint has riot lbeer shown to cause excessive roll
reversals.
6
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DN No.: 1.4-4-7	
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Page: 15
It is apparent that, at this point, no good case exists for controlling
the arrival location on the R -V line. Control can be exercised over
the flyback trajectories, i.e. making them neighboring, within the
linear range of the fuel dissipation phase inertial attitudes.
Figure (5) illustrates what has recently been achieved by proper
selection on the inertial attitude (Reference (D)). Note that a
merging of the flyback trajectories does not make the 14ECO conditions
the same. For the cases shown in Figure (5) the range variation is
from 289.5 to 311.7 HM.
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Pageo.171.4-4-1
ii
5.0 CONCLUSIONS
The following conclusions can be made concerning the use of a 	
f;
constant inertial attitude during the RTLS fuel dissipation phase:
1. The value of the constant inertial attitude can be chosen
E.
either to influence the point of arrival on the R-V line
or to shape the flyback trajectory. Neighboring flyback
trajectories appear significant from the monitoring and
reversion to manual backup aspects.
2. In either case, the control over the parameter variation
is greatest for early aborts and decreases to zero at the
RTLS/AOA boundary.
3. Control over the range to landing site and/or dynamic
pressure at external tank separation, if desirable to
produce more benign entry conditions, can be more
effectively achieved by varying target conditions
(translating the R-V line and changing altitude respectively)
than by using the R-V line arrival point control available.
This leaves the inertial attitude selection free for merging
the flyback trajectories.
I
DN No.. 1.4-4-7
Page: 13
6.0 REFERENCES
(A) FM41 (75-32), Return-to-Launch-Site (RTLS) Preliminary Combined
Guidance and Targeting Formulation Presented to the Powered
Flight Working Group (,April 2-3, 1975), April 23, 1975.
(B) User's Guide for the Space Vehicle Dynamics Simulation (SVDS)
Program Revision 2, JSC Internal Note No. 73-Ft,1-67 November
14, 1974.
(C) "Powered Flight Guidance Ascent Supervisor", John P. Higgins,
December 18, 1974.
(D) t1DTSCO Design Note No. 1.4-4-14, "Return-to-launch-Site
Trajectory Shaping", October 17, 1975.
f


Return-to-launch-site three degree of freedom analysis, constant inertial attitude during the fuel dissipation phase

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