An Entry Flight Controls Analysis for a Reusable Launch Vehicle

14 Pages·2000·0.95 MB·English

by  Calhoun, Phili

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AIAA 2000-1046
An Entry Flight Controls Analysis
for a Reusable Launch Vehicle
Philip Calhoun
NASA Langley Research Center
Hampton, VA 23681
38th Aerospace Sciences
Meeting & Exhibit
10-13 January 2000 / Reno, NV
For permission to copy or republish, contact the American Institue of Aeronautics and Astronautics
1801 Alexander Bell Drive, Suite 500, Reston, VA 20191
I
I
AIAA-2000-1046
AN ENTRY FLIGHT CONTROLS ANALYSIS FOR A REUSABLE LAUNCH VEHICLE
Philip Calhoun
NASA Langley Research Center
Hampton, VA 23681
ABSTRACT Cll3 8Cih513,deg-I
5Cm/5Ot,deg-I
Cm a
The NASA Langley Research Center has been per- 6C,/_, deg -I
forming studies to address the feasibility of various sin- Center of Gravity
gle-stage toorbit concepts for use byNASA and the com- DOF Degrees of Freedom
mercial launch industry to provide alower cost access to LaRC Langley Research Center
space. Some work on the conceptual design of a typical LCDP Lateral Control Departure Parameter
lifting body concept vehicle, designated VentureStar TM LMSW Lockheed Martin Skunk Works
has been conducted in cooperation with the Lockheed LOX Liquid Oxygen
Martin Skunk Works. This paper will address the results RCS Reaction Control System
of a preliminary flight controls assessment of this vehi- TAEM Terminal Area Energy Management
cle concept during the atmospheric entry phase of flight, Angle of Attack, deg
The work includes control analysis from hypersonic flight Sideslip Angle, deg
at the atmospheric entry through supersonic speeds to
final approach and landing at subsonic conditions. The INTRODUCTION
requirements of the flight control effectors are determined
over the full range of entry vehicle Mach number condi- The NASA Langley Research Center, in coopera-
tions. The analysis was performed for a typical maxi- tion with Lockheed Martin Skunk Works (LMSW), par-
mum crossrange entry trajectory utilizing angle of attack ticipated in a conceptual design study for a single stage-
to limit entry heating and providing for energy manage- to- orbit reusable vehicle, lThis reusable vehicle concept,
ment, and bank angle to modulation of the lift vector to proposed byLMSW, isa lifting body configuration called
provide downrange and crossrange capability to fly the the VentureStar TM. This paper documents a flight con-
vehicle to a specified landing site. Sensitivity of the ve- trols analysis performed at LaRC to address the flyabili-
hicle open and closed loop characteristics to CG loca- ty of an early VentureStar TM configuration similar to the
tion, control surface mixing strategy and wind gusts are X-33 configuration. The VentureStar TM concept vehicle
included in the results. An alternative control surface had undergone several design and configuration chang-
mixing strategy utilizing areverse aileron technique dem- es during early conceptual studies. Figure I shows atyp-
onstrated a significant reduction in RCS torque and fuel ical early VentureStar TM vehicle configuration similar to
required to perform bank maneuvers during entry. The the one that will be addressed in this paper. The control
results of the control analysis revealed challenges for an
early vehicle configuration in the areas of hypersonic
pitch trim and subsonic longitudinal controllability.
NOMENCLATURE
rolling moment aerodynamic coefficient
C I
Cm pitching moment aerodynamic coefficient
C. yawing moment aerodynamic coefficient
"Research Engineer, Vehicle Analysis Branch, Aerospace
Systems Concepts and Analysis Competency,
Copyright © 2000 American Institute of Aeronautics and
Astronautics, Inc. No copyright isasserted inthe United States
under Title 17,U.S. Code. The U.S. Government has aroyalty-
free license to exercise all rights under the copyright claimed
herein forGovernmental purposes. Allother rights arereserved
by the copyright owner. Figure 1. Typical Early VentureStar Configuration.
TM
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surfaceesv,idenfrtomtheconfiguratiaofntview,include
An analysis of the entry and landing phases was
bodyflapsattheoutboarednginleocationsin,boaradnd
completed to assess the configuration design flyability
outboaredlevonosnthecantefdins,andtworudderosn
and requirements for the RCS and control surface rates
theverticatal ilsT. hisconfiguratioisnsimilatrotheone
and deflections. An aerodynamics database, representa-
analyzeidnthiscontrolasnalysiwsiththeexceptioonf
tive of the early VentureStar TM configuration, was pro-
thereplacemeonfttwinverticatlailswithasingleverti-
vided byLMSW for stability and controls analysis. The
caltailandrudderforyawcontrolA. combinatioonf
database includes longitudinal and lateral/directional
theseaerosurfaceasndRCS.jetswasutilizedtocontrol
force and moment coefficients for typical flight condi-
andstabiliztehevehicleduringentryandlandingphas-
tions over the entire Mach range. Data was provided at a
esofflight.TheRCS
.jets used for entry control are lo- moment reference center along the vehicle centerline at
cated on the aft outboard surfaces and oriented to fire 66.0 percent of the vehicle reference length. Aerodynam-
vertically for pitch and roll and aft and outboard for yaw ic control surface deflection requirements for pitch trim
control. were determined for the maximum crossrange trajecto-
ry. The surface deflection requirements for trim for this
ENTRY FLIGHT CONTROL OVERVIEW trajectory are shown in Figure 3. This plot represents
equal angle deflections of the body flaps, inboard and
During the hypersonic entry phase, the guidance outboard elevons required to provide pitch trim along
system will command the vehicle to maintain the angle the trajectory from entry to landing. For this analysis,
of attack to limit the maximum heat rate.-" During the the body flaps and elevons deflect equally. The figure
entry phase, bank reversals of up to -+90 deg are com- shows the required deflections with the data adjusted to
manded to adjust the nominal trajectory downrange and CG locations at several positions. ACG of 71.8% of the
crossrange. Bank reversals and angle of attack maneu- reference length isthe aft most location that allows pitch
vers, below about Mach 15,are used to adjust the vehi- trim tobe possible given the assumed maximum control
cle trajectory as required to meet specific flight condi- surface deflections of +_30deg for the etevons and +30,
tions at the Terminal Area Energy Management (TAEM) -15 deg for the body flaps. However, typical thermal
interface at about Mach 2.5. The vehicle iscommanded heat loads on the control surfaces limit the sustained
to follow a heading alignment cylinder and maintain surface deflections to +15deg during hypersonic flight.
flight path angle commands to control the trajectory from In addition, the control deflections for trim are typically
the TAEM interface toapproach and landing. Flight con- limited to _15 deg to maintain vehicle controllability
ditions for a typical entry were provided byaLaRC flight over all flight conditions without excessive use of RCS.
mechanics team, using the Program to Optimize Simu- This is particularly true if the vehicle is either neutrally
lated Trajectories (POST) trajectory analysis tool .3Nom- stable or marginally unstable in either the longitudinal
inal values of angle of attack, bank angle and dynamic or lateral directional modes. The CG should be no fur-
pressure variations with Mach number during the entry ther aft than 70.3% of the body reference length to meet
and landing phases are shown in Figure 2. They repre- these trim constraints for the trajectory flight conditions
sent typical values for an 800 mile maximum crossrange shown in Figure 2. The trim deflections are also shown
trajectory for a CG location of 67.3% representing the sensitivity
60 .....
ot 25
aAtlnaQckle 20 .......... i............. -......... - ....... i............. 20
0
15
90 ..... t
80 ............ .................. .............. _............. -............... 10
Body flap,
Bank 60 ............................. _................... .................... :..........................
angle, 40 ............!................. !...... _' " "
deg 0
cieg -2020 .........."...............
o,o,o o.lA...;..;.:....................i.:......:.... ]
-10
press., [rJ \ _ .....2.............±......................... -t5
Ibf/ft21001 _ J
-20
0 5 10 15 20 '25 30 0 10 15 20 25 30
Mach number Mach number
Figure 2 Maximum Crossrange Trajectory Figure 3 Body Flap, Elevon Deflections for Trim
2
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to
aforward CG shift such aswould occur with the pay-
load returned in an extreme forward payload bay loca-
tion. The hypersonic pitch trim issue was a challenge Static
margin,
for this configuration. For the purposes of this study, it %
was assumed that as the design matured the combina-
tion of CG location, aerodynamic moment center, and
control surface effectiveness will allow pitch trim of the
vehicle within the feasible range of control surface de- 8 o :
flections. 6.0....... ..........._..................i...o............o..-o.-.....o.......ii......................
Time
todbl
long, 4 o i.................. i..................6 0o ""! ...................i...................
VEHICLE.. OPEN LOOP CHARACTERISTICS $ec o o .i............._................i.................!................i...............
i 'ao i i ; : 0 6iOo
Consideration of the vehicle open loop stability char-
5 10 15 20 25 30
acteristics provides some insight into vehicle controlla- Mach number
bility prior to the design of a closed loop system. Stabil-
ity derivatives and indicators are shown in Figures 4-6 Figure 5. Longitudinal Stability Indicators.
for both the aft (70.3%) and forward (67.3%) CG loca-
tions. Figure 4 shows plots of the aerodynamic stability the unstable flight conditions. The relatively large time
derivatives providing an indication of the static stability to double above Mach 3 indicates that closed loop sta-
of the vehicle. The values were computed at the trim bility augmentation can be accomplished with moderate
conditions for selected points along the nominal maxi- demands on control surface rates and deflections during
mum crossrange trajectory shown in Figures 2 and 3. hypersonic and high supersonic flight. Below Mach 3
The derivative of pitching moment coefficient with re- the doubling time becomes shorter reaching a minimum.
spect to angle of attack (Cm)shows the longitudinal pitch during subsonic flight conditions. This indicates moder-
axis mode isslightly unstab°ie above Mach 15and slightly ately unstable subsonic flight conditions for the aft CG
stable down toMach 4for the aft CG location of 70.3%. location and will require relatively large control surface
The vehicle becomes increasingly unstable below Mach rates and deflections to stabilize the vehicle pitch dy-
3 particularly at subsonic conditions near landing. The namics. The situation improves somewhat at the forward
static margin, depicting the location of the vehicle CG CG location where the vehicle has good pitch stability
relative to the neutral point is shown in Figure 5 and is over most conditions above transonic speeds. Subsonic
considered a good indicator of pitch axis stability. The flight is still pitch unstable with relatively low doubling
static margin for the aft CG location remains relatively time.
small for flight conditions above Mach 1.2. Below Mach
1.2 the static margin decreases and the vehicle becomes The open loop lateral/directional stability charac-
more unstable, especially as the angle of attack increas- teristics can be approximately determined from the sta-
es at low subsonic landing conditions. Figure 5 shows bility derivatives and indicators shown in Figures 4 and
effect on the doubling time for the short period mode for 6. The derivative of yaw moment coefficient with re-
c
Cn_
dyn.,
deg-1
_.... _70.3Ojo_g:......................i.i..........i.............].....
Cnl] 0_ ] 3x10 3...i.._.._.o. _,0..,.,.o..0............i............I,..............
2k]k ....... i................ _-._ ...........-_............... i......................................
,co_X.,[,I,:g:., ........._.........+................!................!.................I..
0 10 15 20 25 30 0 5 10 15 20 25 30
Mach number Mach number
Figure 4. Stability Derivatives. Figure 6. Lateral/Directional Stability hzdicators.
3
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spedtosidesli(pC.I_), referred toasyaw stiffness, has a and a negative LCDP value. These criteria demonstrat-
negative value throughout the trajectory down to Mach ed good correlation when compared to the spin charac-
1.2 for the aft CG location. This indicates the potential teristics witnessed from wind tunnel and fight tests for
for directional instability in the transonic, supersonic and several operational vehicles. 4Using the criteria discussed
hypersonic regimes. The lateral stability is indicated by in this reference, the configuration with the aft CG loca-
the derivative of roll moment coefficient with respect to tion remains in the mild departure susceptibility region
sideslip (CI.), referred to as the effective dihedral. The down to around Mach 3 where near zero values of C,,
effective dihedral is negative over the entire trajectory dyn and negative LCDP move it close to the moderate
indicating static roll stability. Sincc the directional and departure susceptibility region. The situation improves
lateral motions of the vehicle are coupled, only a dy- with the forward CG location where there is either mild
namic analysis can be used to determine the actual sta- or nosusceptibility to spin departure throughout the en-
bility characteristics. However, other parameters, which tire trajectory.
use these static stability derivatives, have been shown to
be good indicators of the lateral/directional departure and In summary, the open loop analysis shows regions
spin susceptibility characteristics of aircraft. 4The dis- of concern for the vehicle stability, especially in the lon-
cussion in this reference utilizes a combined directional gitudinal dynamics, over transonic and subsonic flight
stability parameter called C,,pdynamic (C,_, dyn) and conditions. The situation improves for the forward ve-
the lateral control departure parameter (LCDP) for char- hicle CG locations but stability concerns still remain in
acterizing the open loop lateral/directional stability. Plots the subsonic flight regime.
of these parameters over the traiectory flight conditions
are shown in Figure 6. CONTROL ANALYSIS METHODOLOGY
C,ly dynamic is defined as, The flight controls study included the design of
candidate control laws and analysis utilizing a non-lin-
C._.ay,,= C,,13cos(a) - I_ C sin(a) (l) ear six-degree of freedom (6DOF) simulation of closed
loop flight conditions. The results include estimates of
and LCDP defined as. RCS moment and aero surface deflections and rates re-
quired to perform required guidance maneuvers in the
LCDP = C.[j- % (2) presence of wind turbulence. Angle of attack, sideslip
and bank angle response to maneuver commands were
also included in the results. First, a set of linearized dy-
Positive values of both C° dynamic and LCDP in- namic equations of motion were derived at selected flight
dicate no susceptibility to spin _eparture. During the high conditions along the trajectory from entry to landing for
angle of attack, hypersonic phase of flight C. ,dyn. is the purpose of control law design. The resulting con-
positive, but LCDP is negative. C, is small relauve to troller design was included in anonlinear 6DOF simula-
t3
tion to study the robustness of the closed loop design
CIt3which when combined with the ratio of yaw to roll
inertia produces a positive value of C,_, dyn. The value using wind gusts applied during guidance maneuvers.
becomes smaller as the angle of attac_ is lowered and Control gains for flight conditions between design points
the dihedral effect becomes less negative. The LCDP were determined using linear interpolation. Angle of at-
parameter remains negative down to transonic speeds tack and bank angle doublet maneuvers were used to
duc to the adverse yaw produced by the ailerons• The exercise the vehicle rate and acceleration conditions re-
LCDP parameter indicates an improvement in spin de- quired by the guidance system. The vehicle rate and ac-
parture susceptability as the vehicle CG is shifted to the celeration requirements used for this study are shown in
Table I for the various flight phases and conditions. The
forward location. There is little change in C,_, dyn ex-
pected during the high angle of attack flight where the
dihedral effect dominates and C,, dyn. becomes less Table 1.Nominal Maneuver Requirements
positive as the CG moves forward.l_'he spin susceptibil-
ity has been categorized into none, mild, moderate or
Mach Pitch Pitch Bank Bank
severe regions based on the values of these stability in-
Rate Accel. Rate Accei.
dicators. 4Positive values of LCDP indicate no spin de-
parture tendency while moderately negative values, es- 25-20 2deg/s ! deg/s 2 5deg/s 1deg/s 2
pecially when combined with a value of negative Cnb, 20-8 2-4 deg/s I-2 deg/s 2 5deg/s 1-2 deg/s 2
dyn., indicate moderate or severe susceptibility. Mild 8-2 4deg/s 2deg/s 2 5deg/s 2 deg/s 2
departure tendencies are indicated by positive C,I3,dyn.
4
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valuersepresetnypt icagluidancmeaneuvreerquirements ENTRY TO TAEM ANALYSIS
forentryvehiclesT.heinertiapropertieasndtheRCS
locationinformatiounsedinthisstudywereapproxi-
The control laws were designed at 30selected flight
matevaluesrepresentatoivfteheVentureStar
TM config- conditions along the entire trajectory from Mach 25 to
uration shown in Figure 1. the TAEM at about Mach 2.5. The flight control over this
region was designed to provide a required vehicle re-
Linearized models of the lateral and longitudinal sponse to typical bank and angle of attack doublet com-
dynamics for selected points along the nominal trajecto- mands. Table 1shows the magnitude of rates and accel-
ry from entry at Mach 25 to landing at subsonic speeds erations of the doublet maneuvers used in the design
were developed for control law design. The control gains process over the Math number range from entry to
were determined using a Linear Quadratic Regulator TAEM. The entry flight control design uses a set of
(LQR) modern control design approach. The LQR de- elevons, body flaps, and 3-axis RCS for control effectors
sign uses full state feedback and is applied to the lateral/ during the entry phase. The control design utilizes both
directional states separately from the longitudinal states. elevons and body flaps for trim and pitch control through-
The states used in the design are angle of attack, pitch out entry, decent, and landing. Since the different roll/
angle, pitch rate, integral of angle of attack, sideslip an- yaw effectiveness ratio of the body flaps and elevons
gle, roll angle, yaw rate, roll rate, integral of sideslip provide a mechanism for independent roll and yaw con-
angle and integral of roll angle. The LQR performance trol, the RCS isaugmented with elevons and body flaps
index weighting matrices are tuned to provide the de- for roll/yaw control above Mach 2.5. Below Mach 2.5 the
sired vehicle response while simultaneously minimiz- rudder becomes effective and is used for yaw control and
ing the control effector requirements. These gains are roll control is provided primarily by the elevons with
determined by tuning the control response in separate some assistance from the body flaps. The elevons and the
linear simulations of the lateral/directional and longitu- rudder were limited to about +30 deg deflections and 30
dinal dynamics. This technique provides an effective de_sec rates. The body flaps were limited to +30 deg and
method to design and analyze the closed loop vehicle -I 5 deg deflection and 20 deg/sec rates. The feedback
responses over the entire trajectory. The approach was control laws were incorporated into the 6DOF simula-
to design the control system to use the available aerody- tion with linear actuator models. Control surface actua-
namic surface control effectiveness before resorting to tors were modeled with a 2nd order, linear dynamic char-
the RCS. The RCS fuel estimates may be used for com- acteristic with a 4 hz bandwidth. Rate and deflection
parative purposes to evaluate the capability and efficien- limits were imposed on the control surface commands.
cy of candidate control surface mixing strategies. This
approach has been automated using the Matlab software Several control surface mixing strategies were in-
controls functions 5allowing rapid control design for a vestigated before settling on a nominal design. The se-
large number of trajectory points. lected nominal design used the body flaps and elevons
with aratio of3:l body flap to elevon deflections for the
A nonlinear 6-DOF simulation, constructed using pitch control. Body flaps and elevons were commanded
the Matlab 6and Simulink 7software tools, was used to independently utililizing the remaining control surface
evaluate the design and determine the required control effectiveness for roll/yaw control. This approach was
effector requirements. The model includes non-linear chosen since the body flaps had significantly more pitch
rigid body dynamics, 6 DOF aero, Dryden wind gust effectiveness and only slightly more roll/yaw effective-
model and linear actuator response models with com- ness than the elevons. Closed loop vehicle dynamic re-
mand rate limiting. The Dryden gust model is a wind sponse to typical angle of attack and bank angle guid-
turbulence model recommended for study of vehicle re- ance doublet commands for representative flight
sponse to winds for horizontally flying vehicles, sA brief conditions at Math 20 and 8are shown in Figures 7-14.
description of the model is given in the wind gust sec- For the purposes of evaluating the vehicle maneuverabil-
tion at the end of the paper. The 1976 US Standard at- ity the wind turbulence model was not included in these
mosphere was used to generate the atmospheric density doublet responses. The plots show vehicle response,
and speed of sound along the trajectory. This simulation control surface deflections and rates, and RCS torque and
provided an independent evaluation of the lincarized estimated fuel usage from the 6 DOF simulation. The
model assumptions and demonstration of the control bank and angle of attack maneuvers are performed sothat
design performance and robustness. the maximum rates and accelerations in each axis occur
simultaneously. The resulting control surface and RCS
responses represent the total amount needed to perform
these maneuvers at the required rates and accelerations.
American Institute of Aeronautics and Astronautics
56t t _ _ ' I 15• ×105 ,-, i _ ,- _ -.... .... ]
AnQle ._t ,_: ........._................--.--.C.o.mmand J
oi 50[ !7 : '_ _ ! -- Response / I
.... i.......:..t.'.,'.."'_.........-.i......"........
torque,op---' ;_.... r._ ," b........ _'Y.....
441 = _ _ i i
. ._ilil. 221i1i11211112111..iiillliiiiiillli2111
60 _ _
50 ...._.-;.-. _........._..............!............................
Roll,4O ,/ .............-..;..-..-.._..-..
deg 2ot "
30 =_........ ' e.............. _.........
200I
| Roll :.- .... --......
150_- .......... i.... Pitch;........... /'_ ....................
2 t t _
RCS [ [----- Yaw /" _"_"
fuel, 100 ...........:,............... _ .-_-..-..---.- .---
Sideslip,
deg Ibm 50}[ "'" :"_"---'"'----''"-------" ..........
0 10 20 3O 40 50 6O 0 10 20 30 40 50 60
Time, sec Time, sec
Figure 7.Mach 20 Guidance Doublet Maneuvers. Figure 10. Mach 20 Guidance Doublet Maneuvers.
(CG location = 70.3%) (CG location = 70.3%)
301 * J ' t , /
25/ s_"'"!.....................".................... '_ -- Left / /_"'_'-'-"'---_ .... Command
I ,_ /'_ " ;_ : .... Right A_le 25 _ ................... \\ ...... T .......... Response 1
20] /_-.........t N /r' K......_._ _.........._.......................... attack, _ _ ! ;
Elevons, 15L._I }l"_" "_t_t/ "'-," deg 20 t .......:...............:.\L ±. ....i_ .....................1....
151 . _ . i i ; .......... _. .
deg 1:IW;::::(.((:)_:)<[ [ 1".',_'_'!_"::.[:_:::i[[ii:.i..[[i.ii [i.
Roll, so " •.......!.............i i
deg 4200 ,' -----•/!......i..............................
10
151 _ .! ! I
102[ ........_......t...._.... ,_[,--........i.."....,.. _-.........,,1_-.........................i_............................
rates 0$.----tt_-"_z.:.z.r___.TL_l X _._1'. ">'.:.:-:.:_--_ si_s,_, t .........: :......! ....i...........I.............
deg/sec _[ I/ : I ;._'_ II -'- t/] .':............
-5 ,_:-..........;........
_e_: "..../.I..............
-I51 ' _ , _ " , ;
0 10 20 30 40 50 60 0 10 20 30 40 50 60
Time, sec Time,sec
Figure 8. Mach 20 Guidance Doublet Maneuvers. Figure 11. Mach 8 Guidance Maneuvers.
(CG location = 70.3%) (CG location = 70.3 %)
30 r , , , 3°I A: L ! --'Left I
2:5rI°tr_t!: ...... j;_:".................7............ An..l_...i...--L.e.ft... ..............2o_ /\ [ .......,./!'-,-.--,i, . :' /\/_ ! - .... R_h...t......
,,aps,o,_'m°d'I_'"-_--_\_;-,-f_,--:-I__-_:_-, ...........
01 '; i , _ i _' i i
30L " ! " " ! ',._.I.,.i...,.',.. :,'_ ...._..........................
20__=___ _..4......_......................_..........._.......-....
;-_ ,i¸S ,,! i
Flap 10_i,J--'l-'_, ,_.i..--.----I, J'__JI,'--'_/L-_.-n-.,r,. _.................i.....................
rates, 0 - i rates, 0 _ _] [_'-'_ ,_ I,C..._..._-I,'|| ).-_
._o _......_...........i...•.."........_.................i................................
-10 : , !
0 10 20 30 40 5O 6O 0 10 20 3O 4O 50 6O
Time, sec Time, sec
Figure 9.Mach 20 Guidance Doublet Maneuvers. Figure 12. Mach 8 Guidance Maneuvers.
(CG location = 70.3%) (CG location = 70.3 %)
6
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25 t ! r t t tions meet the required maneuver rates and demonstrate
! . _ _ Left
2o_ ,-;: .......... l_ ..........:........ .,",',. .... Ri0ht excellent response to simultaneous pitch and roll dou-
Body _sh i_,.;...J.llk,--==_J';...:..:.................b.le.t.c.o.m.m.an.ds. The angle of attack response is well
tlaps,
deg damped with small overshoot and steady state error.
Good bank turn coordination is accomplished with the
0 induced vehicle sideslip remaining less than 1.5 deg
throughout the bank maneuver.
•"r ! "1 T
10 Figures 8 and 12 show the left (solid) and right
Flap (dashed) elevon deflections and rates required to perform
rates, 0
deg/sec the maneuvers. The elevon deflections nearly reach their
-10
design constraints of 30 deg. with some allowance left
-20 over for response to wind turbulence. The elevons are
0 10 20 3O 4O 5O 60
Time, sec commanded asymmetrically providing mostly roll/yaw
effectiveness. The body flap deflections are shown in
Figure 13. Mach 8 Guidance Maneuvers. Figures 9and 13.The flap deflections are pushed tonear
(CG location = 70.3 %) their maximum positive values during the maneuvers for
the Mach 20 condition. The RCS torque required tocom-
x 105
t.5 plete the maneuver is shown in Figure 10along with an
1.0 estimate of fuel usagc. Low dynamic pressure during this
RCS .5 phase limits the ability of the control surfaces to accom-
torque, --.....!..... !...... j
fNb 0 plish the design maneuver without significant RCS
-.5 torque. The required RCS torque in all three axis may
-1.0 produce excessive requirements on RCS jets for the hy-
personic flight conditions. Guidance strategy may have
80 to be designed to perform required bank maneuvers at
; i ,"_...... "7-
60 ........... ................ _................. !,.:._:j'... !................. i................... higher dynamic pressures near conditions less than Math
RCS I0-15 where the required RCS torque isat areduced lev-
fuel, 40
Ibm el. The pitch axis RCS torque requirements, shown in
2O
Figure 14, are reduced to negligible levels by Mach 8
conditions since the control surfaces are adequate topro-
lO 20 30 40 50 60
Time, sec vide pitch maneuverability. However, the roll and yaw
torques are still substantial, particularly the yaw torque.
Figure 14. Mach 8 Guidance Maneuvers. Further work to tune the control mixing strategy, utiliz-
(CG location = 70.3 %) ing the additional available body flap deflections at Mach
8 may reduce the required roll and yaw RCS torque.
The estimate of RCS fuel usage is computed by in-
tegration of the required RCS force and not the result of A control strategy known as the reverse aileron tech-
acomplete RCS jet model. A constant specific impulse nique was studied as an alternative that could reduce the
of 400 sec was used to represent a value theoretically RCS torque and fuel consumption requirements. This
achievable for LOX/Hydrogen reaction jets at near vac- method commands the aileron ina reverse direction roll-
uum conditions. The resulting fuel values would have to ing the vehicle a small amount in the direction opposite
be scaled to account for different ISP values resulting the intended bank maneuver. This induces a small side-
from atmospheric pressure losses and different fuel/ox- slip angle which allows the large negative dihedral ef-
idizer combinations. The estimates included in this pa- fect to roll the vehicle back toward the direction of the
per were considered to be informative for relative com- maneuver. The sideslip is reduced to zero as the vehicle
parison purposes. Actual vehicle fuel budget calculations roll reaches the commanded maneuver rate. The advan-
would require more accurate RCS modeling. tageous use of the aerodynamic effects of sideslip angle
reduces the roll and yaw control authority needed tocom-
Figures 7and II show angle of attack and roll guid- plete the maneuver. Figure 15 shows the guidance ma-
ance commands and vehicle response for the Mach 20 neuver response utilizing this technique. The roll re-
and 8 respectively. The guidance commands are shown sponse is initiated with a 2 deg roll opposite the
indashed lines and the responses in the solid lines. The commanded maneuver direction which induces about a
angle of attack and roll responses for both flight condi- 1deg sideslip. This is apparent in the response, particu-
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t,;L.....'...............2.............:...........,. i --'Le.
[ ,_ .... Command
| i : I_ _ _ _ .... Right
Body 14 :........... ' i................... _................
20 ........................................... i ............................... flaps, 12 :"-:-'.;.....!............
i5 [ : i ::= deg
50
.ot .// .........\ ......:...............................;............................ 61 i i i i i
°°g 2o iiii21121;1_2./11111111; ....
2o_ ,. _ _ r 1
10 I i i ! L i
Sideslip, deglsec
deg rates,Flap1_210000 i:........ 'i.................... _...................;.:.............
0 I0 20 3O 40 50 60 0 10 20 30 40 50 60
Time, sec Time,sec
Figure 15. Mach 8 Reverse Aileron Maneuvers. Figure 17. Mach 8 Reverse Aileron Maneuvers
(CG location = 70.3 %) (CG location -- 70.3 %)
20 r • ! _ ! , 6/x104 __ i--.°,, I
I ; | " -- Left
181" ";........_...............i-....... ] _ rt"_ Right
....
16 _ | ...... i.......Jl it I ,.....................................
RCS
torque,
deg 12_ II.".....'_.-'.,v',,"lk I_.'.,._.:....:......................ft-.lb...
0363..................._........... ""_-.-....................i..... --':- ..........._-'--.-.-.-..-.Y..a.w.P.__chlt
30r_ i _
2oI..i..........i...:._...-.:.-.-._.-_-------
RCS
Erlaetve°sn, 100t__ iL ...... 1. I_t._ IJ' ' ' fuel, t'.I...i.........!..,."."_......................
o ,=O-tto'1f ..............Ib.m..
tOl ,_-_ .............
I .......
-2o _...............i...................... 5 '¢" i
-3o_--I '_ .............
0 10 20 30 40 SO 60 0 I0 20 30 40 50 60
Time, sec ]3me,sec
Figure 16. Mach 8Reverse Aileron Maneuvers. Figure 18. Mach 8 Reverse Aileron Maneuvers.
(CG location = 70.3 %) (CG location = 70.3 %)
larly during the reverse roll used to complete the ma- Both control strategies where implemented in the
neuver. This technique provided an adequate roll re- vehicle 6 DOF non-linear simulation with wind gusts
sponse below Mach 17as the dynamic pressure increased applied using the Dryden Wind turbulence model. The
above 30psf and performed relatively well to the TAEM gust model parameters where selected toproduce winds
interface at Mach 2.5 where the rudder became effec- representative of severe turbulent conditions. 8Figures
tive. However, the roll response using the reverse aile- 19-22 show the comparison of the maximum elevon and
ron may not be adequate to meet the flight conditions at body flap rates and deflections, RCS torque and fuel
the TAEM interface and a conventional control strategy usage for the two design strategies across the flight con-
may have to be implemented as the flight condition ap- ditions over the range from Mach 25 to 3. Elevon and
proaches TAEM. Figures 16and 17show the elevon and body flap commands, saturated at 30 deg/sec and 20 deg/
body flap deflections and rates needed to perform the sec respectively, resulted inmaximum deflections with-
maneuver. The reverse aileron command uses the elevons in the control surface limits throughout the trajectory.
with some minimal deflection of the body flaps. Pitch As demonstrated in the comparison of the RCS require-
control isperformed with the same body flap/elevon mix- ments for thc Math 8 flight condition, the RCS torque
ing used in the nominal approach. A comparison of the and fuel usage were reduced substantially at flight con-
Mach 8 responses shown in Figures 14 and 18demon- ditions along the trajectory less than Mach 15 for roll
strates a significant reduction inroll and yaw RCS torque and Mach 20 for yaw. The reduction in yaw RCS and
and fuel usage for the proposed technique. increase inroll RCS usage above flight conditions around
American Institute of Aeronautics and Astronautics