The results are presented of a study to show the effect on flyback trajectories of constant inertial attitude during the fuel dissipation phase of a return-to-launch-site abort. It is shown that the value of the constant inertial attitude can be chosen to shape the flyback trajectory

Bown, R. L.

English

NASA Technical Reports Server

General Disclaimer 
One or more of the Following Statements may affect this Document 
 
 This document has been reproduced from the best copy furnished by the 
organizational source. It is being released in the interest of making available as 
much information as possible. 
 
 This document may contain data, which exceeds the sheet parameters. It was 
furnished in this condition by the organizational source and is the best copy 
available. 
 
 This document may contain tone-on-tone or color graphs, charts and/or pictures, 
which have been reproduced in black and white. 
 
 This document is paginated as submitted by the original source. 
 
 Portions of this document are not fully legible due to the historical nature of some 
of the material. However, it is the best reproduction available from the original 
submission. 
 
 
 
 
 
 
 
Produced by the NASA Center for Aerospace Information (CASI) 
https://ntrs.nasa.gov/search.jsp?R=19760026147 2020-03-22T13:49:23+00:00Z
.n
MN
1
,Oh
z
.	 -	 •	 ^..nH^IfPEZ
 a^,ry«...rte ...:vw.rn...a, o
	 ^...rY11^... - •fi
•
MCDONNELL DOUGLAS TECHNICAL SERVICES CO.
HOUSTON ASTRONAUTICS DIVISION
SPACE SHUTTLE ENGINEERING AND OPERATIONS SUPPORT
^	 A r
` S
DESIGN NOTE NO. 1.4-4-14
U) ^	 RETURN-TO-LAUNCH-SITE TRAJECTORY SHAPING
q ul	 1:	 s
MISSION PLANNING, MISSION ANALYSIS AND SOFTWARE FORMULATION
Q-
m
U)
17 OCTOBER 1975
APPROVED BY:
W. W. Hinton, Jr.
2o272B29 ^	 FPB Work Package Manager
488-5660, Ext. 240
OCT 1976	 ^,
acova
Np^q S11 p1111Y. ^,
INPUT BRp`	
C:^'
f
APPROVED BY: )tT^ f ^., ?,^ ;
W. E. Hayes
Project Manager
Mission Planning, Mission
Analysis and Software
Formulation
488-5660, Ext. 266
NN
r4
W U
E+ U)
H Ui U
U? b
::
1
 CP V
U =1 if`
O
A: 1 tfi
r4
1'-4 U
O a) x
H C
1 ^ a
x o
U n7
E- 4 f
^, U1
7- (ll
01 a •ri
O Kc >
O ^ H
P to (L
111 U]
P• >4
1^ ucti
U E- U
1 U 'rl
^ ^ U
This Design Note is Submitted to NASA Under Task Order
No. D0303, Task Assignment 1.4-4-D in Fulfillment of
Contract MAS 9-13970.
PREPARED BY: /^12A^	 APPROVED BY: `,
R. L. Bown	 L. C. Winans
Senior Engineer	 Task Manager
488-5660, Ext. 243	 488-5660, Ext. 243
rs{^
1•
T
4
4
1.0 SUMMARY
This document presents the results of a study to show the effect
on flyback trajectories of constant inertial attitude during the
fuel dissipation phase of a Return-to-Launch-Site (RTLS) abort.
Results are presented which show that the value of the constant
inertial attitude can be chosen to shape the flyback trajectory.
i	 5.
DN No.: 1.4 -4-14
Page: 2
2.0 INTRODUCTION
Preliminary RTLS guidance and targeting software for the Space
Shuttle is documented in Reference (A). This note documents
another in a series of performance verification studies planned
to verify the adequacy of that software.
After a space shuttle main engine (SSME) shutdown and crew
selection of the RTLS mode, a fuel dissipation phase will be
required. The use of a constant inertial attitude during this
phase to shape the trajectory was proposed in Reference (B).
This study, was conducted to determine the feasibility of shaping
all RTLS trajectories neighboring to that of a mode boundary
abort.
p
.aa!^y	 F ...	 }	 .......,.	 ....	 ..	 :fx,	 ..., _.., ws.,r..	 in.,...•	 ,F..,.,,^,,,,..a	 w...,o.m,^.	 ..	 ...,..	 ...	 ._ v
	..	 ^.	 ..	
..
DN No.: 1.4-4-14
Page: 3
3.0 DISCUSSION
The abort conditions or times used in this study were liftoff,
Solid Rocket Booster (SRB) separation, 140, 180, and 220 seconds
from liftoff. The trajectories were shaped in the inertial
velocity - altitude plane to obtain a maximum altitude of approximately
385,000 feet similar to that of the mode boundary abort.
The constant inertial attitude (zero body rates) was maintained
by holding a constant inertial thrust direction. For the three
degree of freedom simulation used in this study the thrust axis was
alined with the X-body axis. The inertial thrust angle is
measured from the launch site local horizontal plane which was
fixed in Earth Centered Inertial coordinates at time of launch.
The thrust angle is measured positive upward in the nominal down-
range flight configuration shown in Figure 1.
The excess Orbit Maneuvering System (OMS) and Reaction Control
System (RCS) fuel was dissipated by igniting the two OMS and the
four aft axial RCS engines. Subsequently a preselected quantity
of OMS fuel was burned by the same RCS engines to insure complete
consumption of the ONIS fuel before main engine cutoff (MECC). The
guidance software computed the time for turnaround. The RTLS abort
procedures are assumed to commence at time of engine shutdown
except for shutdowns prior to staging when the procedures commence
at staging.
tW--j
R
a
DN No.: 1.4 1-14
Page: 4
W
w ^-o¢O
W J
J^-+V
Q 1-- Z
D W Q
J 9 J
^Z
Q'
v
Q'
2Q
QV
Q
ti
O =Z ~R
I	 ^V W
w
Ov
Q
aJ
o Q
^O
w Z
^ v
a^ o
z
a Q
-'0"0
C!
Qt
C
ro
'C3
r
4J
41
Q
r
ro
L
OC
4-
O
c0T
.r
CrW
D
dL
rn
-7-1777=
DN No.: 1.4-4-14
l	
Page: 5
This study used a three degree of freedom simulation contained on
a modified Space Vehicle Dynamic Simulation (SUDS) 2.3.11 milestone
file (Reference (C)) 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 (D)).
b) Addition of the thrust termination logic (Reference (D)).
The targets input to the Powered Explicit Guidance (PEG) module
were biased to the ;gain Engine Cutoff minus ten seconds (MECO-10)
conditions of 310,000 pounds of total weight, 230,000 feet altitude,
and a four degree 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 = .069VE - 110.1
For thrust termination the target was the RI MECO R-V line:
R = .068VE - 171.5
All ranges are from the landing site at the Western Test Range in
nautical miles and the relative velocity (V E ) is in feet per second.
The inertial attitude required to shape an unperturbed RTLS trajectory
to that of the mode boundary is computed as a function of the inertial
velocity at the time of abort. For aborts occurring prior to staging
the velocity at abort is defined as the inertial velocity at staging.
The algorithm includes partial terms to modify the inertial attitude
for off nominal SRa staging conditions.
'01 ^,F ^:.. ,^
DN No.: 1.4-4-14
Page: 6
6 - 90.0 - IGNOM + EA1/(AhMAX/AONOM)I
	
(3-1)
where
ONOM = nominal inertial attitude
EAh summation of quadratic terms and numerical
partial derivatives to account for off-
nominal SRB staging conditions. In
particular,
EM = [ (AhMAX 
J
/ AVI
sAV  
s
+ 
(nm)  Ah
Ah 
	 s
+ eL)AyI'sJJ
n'S
Ah VAX /AUNOM - change in maximum altitude with respect
to a change in attitude.
s - denotes evaluation at S"B staging.
h - altitude
y - inertial flight path angle
AV 
I'S' 
Ah, Ay  - actual minus nominal conditions at SRB
staging.
.W
DN No.: 1.4-4-14
Page: 7
4.0 RESULTS
The interim analysis leading to a selection of the algorithm started
by obtaining the maximum altitudes for unperturbed RTLS trajectories
at three attitudes as shown in Figure 2. The maximum altitude for
a
!	 the mode boundary abort was used as the parameter to choose the
required attitude for the five study cases. These angles are shown
in Figure 3 plotted versus the inertial velocity at abort. The
data points were curve fit as a quadratic function of inertial
velocity at abort. The shaped trajectories for the unperturbed
cases are shown in Figure 4.
The nominal SRB staging conditions were perturbed in velocity,
flight path angle, altitude, and range. The perturbed trajectories
were generated at the nominal attitude to determine the change in
maximum altitude with respect to the perturbed condition at SRa
staging. The numerical partial derivatives are shown in Figure 5
as a function of the inertial velocity at time of abort. The
maximum altitude is insensitive to a change in range at SRE staging
since the result is that the trajectory is translated in range
only.
The partial derivatives were curve fitted with quadratic functions
of inertial v^-Iocity. The change in altitude caused by each off
nominal SR6 staging condition is obtained by multiply i ng each
partial derivative by the associated difference between actual
and nominal 'W staging condition. The result is a cummulative
toV
m c
O O
Te: OL1
DN No.: 1.4-4-14
Page: 8
O	 O	 G1	 OU	 U
O	 O	 ^ :3 4-+	 O
F- 4' C)O	 Q	 •r •r C)	 O
Ln	
O	 X 4^ 4-	 O
V	 RI ^-	 MQ 1
0
0
o► 	 r_0
r
i
NC
O	 +^
O	 i
coo
	 .nQ
N
fL
b
	
to	 ^
O d 4J
	
0 4-	 •r
C)	 +_3
	
f l, 1
	
r
Q
	
i	 EO^
 E
	Q 	 •r
X
	
4-	 rpO OO
O 41 N
	
f.0 E
	 -^
•r
F-
4-) C
	
to	 O
	
A	 4-)
	
4-3 	 U
	
O •r 	 N
O U 4-
O O 4-
	
1n r-	 W
QJ
	^ 	 fL
.0
	
r	 :3
	16 	 4J
	
•r	 •r
	
4J	 4-1
	
i	 4JO U d
O C
O «	 r-
_7	 r0
•r
i
C/
C
r-.
O	 Q1
O	 tT
O	 C
Cl)	 ed
S-
0n
N
a
S,
:J
V1
LA-
L
m
60
50
Attitude Angle
-Degrees
40
i
i
T
ON No.: 1.4-4-14
Page: 9
i
3000	 4000	 5000	 6000	 7000
	
8000
Inertial Velocity at Time of Abort - fps
Figure 3 - Attitude Angle Required to I•latch Mode Boundary
Trajectory vs Inertial Velocity at Time of
Abort - Unperturbed Trajectory
DN No.: 1.4-4-14
Page. 10
L
`	 N C O O O O O
W Q Ch f^ Q	 COvv...... 1Aal o,
b o0 0 0 C 'O
+s++ Nom.-Ni. C
O++ -
•+
t- W
.0 4j
A'(3k p
v) v
W Q
Q • Q^
• * C
CD W
u e0
Ow
n1 J >,
• + y
tl) W 1- U
S O
111	 1'^	 I '~y
i
L
4W C
cn
vv
O W r
C^
1	 C—
^ N
.
C 
+
1 W	 L
d
t f'	 '.
• 4
-' 11J	 •r
Q1L
in
	
It)
	
^n	 1l	 rn
W	 (r) W	 tv W	 W	 ^^	
►^
N
QJ
.L
O
va
tL1
L
F-
r
ro
Lv
ItEpli uclBUJ'I'Y Oy 7'1'
ORIQLNAL PAGE IS PO( w
mm
m cx
4n	 t/1
d
m
N
t
I
Ld
C
X W
re
a0	 C) F  v)) C^
C)	 C) X- r U O
C) 7 -- p LO
CV	 r r
C3	 C x cn O C C-)
C3	 un ro n • Cj u)
N	 r ( V)
r_
X I fn
C)	 u)	 v	 Cl" c	 ^n
(.	 N	 C)
tv	 .-	 r.	 r. r
c0
4+
^	 b
.v
>o	 aN
N
co
Vf
U^
	 H
^	 CO
w
H ^
a J-
4-	 :3
O ^ yD 4J d
^O L
O C)
.0	 a.+4
4- V
O	 C)
C.
a	 VnO E vO •r ^
U F-
L	 r
♦-+	 +J
T •^
iJ	 Q1
u
o a^+
CD . - •r
O Q) 4J
p } ^-
r
w f-
4-J
	F
QJ	 x
C	 rp
CD	 C
C3	 •r•
M
41
C
b
tV
u)
C)
i
G'.
U.
i
i
DN No.: 1.4-4-14
Dage: 12
change in maximum altitude (Uh) caused by off nominal SRB staging
conditions. The nominal attitude must be modified t,o account for
this Uh. The change in maximum altitude with respect to a change
in the nominal attitude (Ahmax/60nom) was obtained from Figure 2
and is shown in Figure 6 as a function of inertial velocity at
abort. The required change to the iominal attitude is obtained
by dividing 
Uh by hmax/AOnom)'
The resulting algo- :hm is
0 = 90.0 - 0, + ' JIMV,AX) AV	
+ ( Aht" ) Ahs
iti0:^1	 I , s
	
I ^ s	
`	
s
+ I ^) 
"I	
nht± 
l	 (4-1)
I`  AY I s J	 +s(
	
nO NION )J
(^h	
,l
The tern s ONOt1' AV 
I's	
,	
nhtSAX	 Atf^tSAX
' 	 s	
YI,s
and 
M 1 AX	
w , re fitted by quadratics and the coefficients (for the
NOM
Mission 3A simulation used in the study) are shown in Tabl, T.
For example:
0N01! ' .474 + . 0121V I,l^F; - 7.31 X 10-7V1,AE3
:here
VI,A(; - inertial velocity at time of ahort
The conditions at abort attitude, and P-V line arrival points are
shown in Table II for the noriinal study cases. The algorithm has
tested for perturE-,ed SI:L staging cor,ditio!!s of + 200 feet per second
OOO
M
3a N
D a
4-
+-)
i
O
d,
O
O 4-
O O
to
Q)E
•r
H
CO +•)
0 r
In U
O
O
7
t^
rO	 4-)
C-)	 S-
O 4)
^' CN
0Cl
ON No.: 1.4-4-14
Page: 13
CU	 O	
O	 O
O	 O	 O	 O
Cl	 X	 0
(VC)CJ
O	
LC)lf)	 rJ
 
r LE	 '^
a^
v
c
c	 11
ai
•v
a^
r
4J
41
d
C
r
4)
C
f^
sv
O
41
•-)
V
O
CL
LA
4)
sy
•r
S
Q)
41
d
E
3
E
•r
K
C
•r-
4)
Q1
sU
t0
4)
3
C)l
tL
DN No.: 1.4-4-14
Page: 14
TABLE I
QUADRATIC COEFFICIENTS - Mission 3A RTLS Trajectory Shaping
Term
Quadratic Coefficients
a b c
011014 .474 .0121	 • -7.31	 X 10-7
AhNAX/AVI,s
Ah 
VAX 
/Ahs
AhM-AXIAYI,s
AhtAX/AGNOM
450.
3.31
67600
26809
.01023
-6.75 X 10 -4
-14.91
-5.05
6.12 X 10-6
4.36 X 10 8
9.22 X 10 4
2.12 X 10-4
{^, NC
0
t ^COU
rt
C
O
^ Z
k—r
--^ N
00 LUQ r—,
F— cr0
U
W
7
Q
cmz
►-r
x
i
s
cO
i
n.
4!
h
4-J
ti
V)
C
O
7
J
DN No.: 1.4-4-14
Page: 15
_tEP ROD UCIBILITY OF THI,
cia(INAL PAGE I3 POOR
of velocity, + 1.5 degrees of flight path angle, and ± 15,000 feet
of altitude. Single perturbations resulted in A0 values from 2.0
to 4.1 degrees for aborts occurring at 180 seconds and earlier.
Perturbations of the late 220 second trajectory yielded significantly
larger corrections to ©: 5.1 to 10.6 degrees. The corrections
are larger at the late abort due to the limited time available to
affect the trajectory.
For a late abort if the perturbations result in below nominal
performance the A0 will be a correction to loft the trajectory.
This is acceptable since the rotation is in the direction of the
subsequent turnaround maneuver. If_the performance is significantly
better than nominal the A0 will command a rotation toward the earth
to decrease the rate of change of altitude. The rotation is opposite
to that of the turnaround maneuver. To prevent undesirable reverse
pitch maneuvers the algorithm requires a limit on the computed attitude
for late aborts. If the computed attitude is lower on the horizion
than the actual angle at abort, the current actual angle becomes
the commanded angle. The higher altitude that will occur during
flyback must be accepted as a characteristic of late aborts.
The conditions for selected perturbed cases are shown in Table III.
The shaped perturbed trajectories shown in Figure 7 were generated
using the algorithm. The MEC0-RV arrival points varied from 6812
to 7188 feet per second relative velocity and 291.9 to 318.6
N
C
O
4J
C
O
U
'C7
L3)
i
^ N
W 1
J
co Lnd W
CL'
O
U
W
7
H
C'3
Z
S
U
H
DN No.: 1.4-14'-14
Page: 17
A
Qp N O m m NN Op M M et ^ r
O O
t00to %0
O
U
W
t0 CO tO r ct M
n Z O C) O ON mM M M M N N
n M N C) G
CD
O
.
C '^ U) mr M Ln
fL tD t0 N M O N%
C r Ln Ln O) t7 M r-b M co d CO M co
S. O co .- r- O ItO Ln LO ^ OM Ln N.
.0 M M CO -,T r- tO
d +^ M M %0 m co CD
—F- i , M r Cn
+J d LL. r r r N M N
LLs
Ln
C
+J r N N N r N
p	 C'3 n tO r O O Lf)
O .-. LLJ
Cr CO Ln r -;;f O C )
U N N N r r
N N N r N
H d ^^ u v
La. - t0 N f^ CO Cl)
h f,, Ln ^ O fl-
CO Ct 00 00 tO r
•a' LP fl, f\
a
f_
f-
+3
i C1 2 U U U U
o a a a a a, a
n N N V) N V) N
Q
co co C) O C) O
w O' tf' 00 N N
V) V) r r N N
C CO O
•r •r	 ;
+J 4J
i L
dC.
a fL
V) Ln
co C7
ine CC
Ln N
i-) i.)
to b
C C
O O
•r •r
N N
i i
a an a
VL N
•r •r
a a
•r ft3
N ^
O a
Cl- Z=
L L
r N
icli
NC1O%Q
Ln w' M u7
0 0 0 0
ctcoNN
r r N N
Q Z k' X
DN No.: 1.4-4-14
Page: 18
	
in	 in	 to	
U)
	-1 4
	
^ ^
	
cv w	 w
3-on 1 1 ou
+' LPRODUCIBILII'Y OF THE
"MINAL PAGE IS POOR
,g
aw
L
O
Q
m C
^O	 b
w d
7
t
po
^
trl	 d
mw
V
O ^
w iy o
++ui w
cc
4 0^r
r
cn
=;w	 aC
N
7
QJ
1
	vi W
	 41	 V
r ^
	
4	
s-
O
	C 	 ~'U
0. ci
	
V>	 i_F-
4 'YW	 >>L v
o u
	
u	 s-u
a. 9
u
^U
cuL
cn
C7
LL-
DN No.: 1.4-4-14
Page: 19
nautical miles. The results are from additive dispersions, that
is a higher velocity was combined with a higher flight path angle.
The algorithm provided a good result for the double dispersed
cases which are considered less likely to occur. The algorithm
requires storage of 18 prelaunch computed coefficients and
nominal SRB staging conditions, the measurement of three trajectory
states at SRB separation, and the use of the actual inertial
velocity. The computation of the commanded attitude requires the
evaluation of a combination of quadratic terms and limit logic.
a
i
i
r
DN No.: 1.4-4-14
-	 Page: 20
i'
5.0 CONCLUSIONS
The following conclusions can be made concerning shaped RTLS
d
trajectories:
C
1. It is possible to shape the RTLS trajectories within
z	 a few thousand feet altitude band of the mode boundary
abort trajectory maximum altitude.
2. The shaped trajectories arrive within a 425 feet per
second band on the MECO-RV line.
3. The algorithm is simple incurring minimal requirements
for the onboard software and accounts for perturbations
from nominal SRB staging conditions.
4. Near the Mode boundary a limiter needs to be introduced
to avoid pitch reversals.
7q
DN No.: 1.4-4-14
Page: 21
6.0 REFERENCES
(A) FM41 175-32), Return-to-Launch-Site (RTLS) Preliminary Combined
Guidance and Targeting Formulation Presented to the Powered
Flight Working Group April 2-3, 1975), April 28, 1975.
(B) MDTSCO Design Note No. 1.4-4-7 "Return-to-Launch-Site Three
Degree of Freedom Analysis, Constant Inertial Attitude During
the Fuel Dissipation Phase", October 17, 1975.
(C) User's Guide for the Space Vehicle Dynamics Simulation (SUDS)
Program Revision 2, JSC Internal Note No. 73-FM-67 November
14, 1974.
(D) "Powered Flight Guidance Ascent Supervisor", John P. Higgins,
December 13, 1974.


Return-to-launch-site trajectory shaping

Abstract

Similar works

Full text

Available Versions

NASA Technical Reports Server