NASA Technical Reports Server (NTRS) 19960016113: Airbreathing hypersonic vehicle design and analysis methods

14 Pages·0.87 MB·English

by  NASA Technical Reports Server (NTRS)

NASA-TM-111249
AIAA96-0381
AirbreathingHypersonicVehicleDesign
andAnalysisMethods
MaryKaeLockwood
DennisH. Petley
JamesL.Hunt
NASALangleyResearchCenter
Hampton,Virginia
JohnG.Martin
LockheedMartin
Hampton,Virginia
WorkperformedunderNASAcontract
34thAerospaceSciences
Meetingand Exhibit
January 15-18,1996/Reno, NV
Forpermission tocopy orrepublish, contact theAmerican Institute of Aeronautics andAstronautics
370 L'Enfant Promenade, SW •Washington, DC 20024

AIRBREATHING HYPERSONIC VEHICLE
DESIGN AND ANALYSIS METHODS
Mary Kae Lockwood, Dennis H. Petley, and James L. Hunt:
NASA Langley Research Center
and John G. Martin: Lockheed Martin
ABSTRACT methods in each of the disciplines. In the center of
the figure, HOLIST isbeing developed as a working
The design, analysis, and optimization of airbreathing environment for design, analysis, and optimization
hypersonic vehicles requires analyses involving many of airbreathing hypersonic vehicles. The basic syn-
highly coupled disciplines at levels ofaccuracy thesis system in HOLIST iscurrently operational.
exceeding those traditionally considered in aconcep- As itisfurther developed, HOLIST will include ele-
tual or preliminary-level design. Discipline analysis ments from all of the disciplines, with upgrades con-
methods including propulsion, structures, thermal tinually being made as discipline methods advance.
management, geometry, aerodynamics, performance, Following is a brief introduction tothe airbreathing
synthesis, sizing, closure, and cost are discussed. Also, hypersonic vehicle desigla process, discussion of
the on-going integration of these methods into awork- selected discipline methods and the current and
ing environment, known asHOLIST, isdescribed. planned capabilities of HOLIST.
INTRODUCTION
The Systems Analysis Office (SAO/Hypersonic
Vehicle Office) atNASA Langley Research Center
provides evaluation, analysis and design of hypersonic
airbreathing vehicles for both industry and govern-
ment. A wide range of vehicles and missions are inves-
tigated, including single-, two-, and three-stage-to-orbit
vehicles, aswell as endoatmospheric cruise and accel-
erator vehicles. Due tothe highly integrated engine/air-
frame and the extensive flight envelop inherent inair-
breathing hypersonic vehicle design, analyses involve
many interdependent disciplines with high sensitivities
among the design variables and ahighly nonlinear path
design spaceL It istherefore necessary toresolve air-
breathing hypersonic vehicles toa preliminary design Figure 1. SAO Hypersonic Airbreathing Vehicle
level, even for what would traditionally be considered Analysis, Design and Optimization.
as conceptual design. With this amount of detail
required as well as the requirement for a short response
time, analysis methods have been developed and HYPERSONIC VEHICLE DESIGN/
improved toprovide both rapid and accurate results. ANALYSIS METHODS
This paper describes the advancement inSAO design
and analysis methods during the past six years. A schematic of the design/analysis process is
shown in Figure 2. The process begins with a
Figure 1illustrates the set-up of the Systems vehicle geometry definition, as shown in the left
Analysis Office, with technical experts and analysis of the figure. Propulsion, aerothermal and trajecto-
*Copyright © 1996 bythe American Institute ofAeronautics and Astronautics, Inc. No copyright isasserted in the
United States under Title 17,U.S. Code. The U.S. Government has a royalty-free license toexercise all rights under
the copyright claimed herein for Governmental purposes. All other rights are reserved bythe copyright owner.
ryanalyseasre completed to yield the propellant tion of the forebody flowfield properties and the
fraction required (PFR). The propellant fraction mass capture are also critical to resolving the net
available (PFA) is determined from packaging, thrust. Therefore, the ramjet/scramjet cycle code,
structural and thermal management analysis, as SRGULL 2,developed primarily in-house, uses a 2-D
well as weights prediction. In the upper right of Euler calculation on the forebody and inlet, coupled
the figure, the plot of PFR and PFA versus TOGW with aboundary layer solution, to predict the fore-
illustrates the process of closing a vehicle. For body/inlet drag and the flow properties entering the
example, if upon first analyzing a vehicle, the PFA engine. The ramjet/scramjet solution isthen complet-
is less than the PFR, the vehicle must be sized up ed using a 1-D cycle analysis with equilibrium chem-
to a higher TOGW until the curves intersect, thus istry and multiple steps through the combustor.
closing the vehicle. Note, however, that to achieve Finally, the nozzle forces are resolved using the 2-D
the accuracy required for airbreathing hypersonic Euler and boundary layer codes. A 3-D Euler capa-
vehicle design, the vehicle is closed on volume bility isnow being implemented into the code.
and area, in addition to weight.
( SEAGULL (2DEuler) SCRAM (1D)with EQChemistry
. . ,. |* Fccebody/]net shock osses °Combustorcycleanalysis
lnlegrale(u I {Control_Ju_ process)
automated
Costing into oneoperation HUD(Boundary Layer) SEAGULL (2D Euler)
SRGULL Io Forebody/]ntet/comT0ustor/nozzle ° Nozzleexpansion losses
_oHeatandfdctlonlosses
___
Trajectory t -- "N
[p_pell_atnaflfyascius_ required) s:
Vehicle geometry
propulsion design
structural design
packaging Input: - Geometry Capabilities - Laminar/transitionallturbulent boundarylayers
- Boundary now include: LOCalflowvector
conditions - Engine flowfield vehicle trim (2D pitch)
- Fuelschedule Lift,thrust, moments
-Thermal balance
Upgrade: -2D/3D Eu]er - RAMStability Model
- Isolator pedo nmance
* LOXaugmentation
Figure 3. Tip-to-Tail Scramjet/Ramjet Cycle
Analysis, SRGULL.
Figure 2. Vehicle Design/Analysis Process in
SAO. Capabilities in the SRGULL code include the
analysis of laminar, transitional and turbulent
Each of the disciplines shown in Figure 2 are criti- boundary layers; engine flowpath forces such as
cal to the design of an airbreathing hypersonic lift, thrust and moments; and LOX augmentation.
vehicle. However some disciplines are more tradi- To first order, a thermal balance can also be accom-
tional in that they may be found in other speed plished. Given the wall temperature, heat flux to the
regime analyses. The three shaded disciplines, walls (calculated by the code) and the fuel injection
propulsion, thermal management and structural temperature, the amount of fuel required to actively
analysis, are unique to airbreathing hypersonic cool the vehicle is determined. This fuel flow rate is
vehicle design. As a result, SAO has developed then used to predict the net thrust for a thermally
unique tools and capabilities in these areas. balanced system. Particularly at high hypersonic
flight Mach numbers, the increased fuel flow rate,
Propulsion which isgenerally above an equivalence ratio of
one, can significantly increase thrust. The predic-
Airbreathing hypersonic vehicles are characterized tion of coolant fuel flow rate is further refined in
by highly integrated engine/airframes as illustrated in the thermal management analysis as described in
Figure 3. Since the net propulsive thrust of an air- the corresponding section below.
breathing hypersonic vehicle isa small difference
between two large forces, namely the combustor/ SRGULL 3also has the capability to predict
nozzle thrust and the forebody/inlet drag, it isneces- engine unstart, which is another unique feature of
sary to resolve these forces accurately. The predic- this cycle code. Figure 4 shows the isolator/ram-
2
jet/scramjekteel-lineatthetop.Thearrowsmark 5O
pointswherefuelcanbeinjectedT. hefourplots
45
showthepressuredistributionthroughtheengine
40
asafunctionofdistancealongtheengineforvar-
iousfreestreamMachnumberswheretransition
3035 Shoulder --_
betweenpureramjetandpurescramjeotccurs.
pip1 25 i z :
Notethatinthetopplot,fuelisbeinginjected
* Cowl side" "1.Experimental _i z ::
fromthemiddleinjectorsatanequivalencreatio 20 Ix Body side J data i * /r_ ;;
of.3andfromthedownstreaminjectorsatan 15 ..... SRGULL i*_,
;H _--"Expansion
equivalencreatioof.7.Alsonotetheriseinpres-
surethatoccursupstreamoftheqb=.f3uelinjec- 10 _'
5 b.
tor.If morefuelweretobeaddedatthisfuel 0 I I I I _1 I I t
Distance
injectorthepressurerisewouldbepushedfarther
andfartherupstreamu,ntilatsomepointan Figure 5. Isolator Model Comparison with
engineunstartoccursN. otethatasthefreestream Mach 4 Experimental Data 4
Machnumberincreasesth,efuelcanbeinjected
fartherupstreamwithoutcausingthedisturbance The Concept Demonstrator Engine (CDE) is cur-
tomoveupstream. rently being tested in the 8' diameter hypersonic
tunnel at Langley. SRGULL has also accurately
predicted the pressure distribution, including the
pressure-rise magnitude and location, as compared
i I
I i
to the experimental results.
Pressuredistri_ ; i
Structures
2_- II //i" ""x. Ramjet
atM 5 / i / i
oo= J / / I ] X.
Hypersonic vehicle structures are characterized by
O/ I , ', ! , --
,, ,=% % thermal loads that are as high asthe mechanical loads.
6 f "
f', 1 Again, due tothe design sensitivities inherent inair-
41- I .jk
atMoo=5.5 _/', ./ ',', breathing hypersonic vehicles, itisnecessary toaccu-
el- j./ , x.. rately predict structural weight, as well as the aerother-
k moelastic flight response ofthe vehicle even atthe
P'l-h 4 *=_.7 '.3
fJ ,,}x., conceptual/preliminary design level. Some of the
codes used inthe Systems Analysis Office include
I 1 / I \]
atM==6 2I/-- /// Ii \",l_ Pro/ENGINEER, for CAD (SAO iscurrently switch-
/ E / I IX. ing over tothis code from another CAD package);
1 II--4/r II II"...
MSC/NASTRAN, P3 PATRAN and RASNA for
O- [ , A' I
T
6 F (_-i 1.0 finite element analysis topredict element loads; and an
4f FE ___I Scramjet in-house developed software package, ST-SIZE 5,to
atMoo=7 /_--- ix
2_-//i IIX", perform panel failure mode analysis and panel sizing.
Figure 6 shows a schematic of how a structural panel is
Figure 4. Ramjet to Scramjet Mode Transition sized. Starting onthe left-hand side ofthe figure, initial
with SRGULL 3. element stiffnesses, thennal coefficients, thermal and
mechanical loads, and the finite element geometry are
input into the finite element analysis code. Forces on
Figure 5 shows an experimenP run in a Langley
each of the elements are then determined. Moving to
tunnel to study the effects of geometry changes
the right ofthe figure, the element forces, material
on isolator flowfield characteristics. As shown,
SRGULL accurately predicts the pressure distur- selections and panel and beam concepts are input tothe
bance in the isolator. ST-SIZE cede. Here upto30failure mode analyses in
strengtahnd26failuremodea_lysesinstabilitayre mesh. The resulting panel weights are shown inthe
performeadn,dthepaneislsizedtomeetthesfeailure lower left-hand comer. Using the enhanced ST-SIZE
modesG.iventhenewpanedlesign, the dement stiff- code, withitscorrec_on terms for membrane-ben_g
nesses and thermal coefficients change and the FEA coupling input into MSC/NASTRAN, the themaal
must recalculate the element forces. This iterative moment isonly 1% different than thatpredicted by the
process confines until convergence isachieved. The fine 3-D subscale-sized mesh. The other element loads
net result isthe minimum panel weight, which results show similar error comparisons. The resulting panel
from a maximally stressed panel that also meets each of weights forthiscalculation are shown inthe lower right-
the failure mode tests, allwithin the margin-0f-safety. hand comer. Note thatthe more accurately predicted
weights aresignificantly different than those for the s_a-
dard 2-D panel calculation. Thus withthe enhanced ST-
ST-Size Automated Iteration
SIZE code itispossible toproduceaccurate structural
Model I_ata - ] [ Design Data 1 weight predictions for airbreathing hypersonic vehicles
__Efloemmeenst V/ --_ J_-7_-- -Jfo'qEple_me_n_t __I II ina rapid pre'lmainary/conceptual level design. This
method alsolends itselftothe loose-coupling ofa FEA
code with an aerothermal CFD code, enabling accurate
predictions ofa vehicle's aerothermodastic response.
Lo 3d q-Nsapproach iscurrently being pursued bySAO.
Elemen_tti"n_"s / fl___o_,l I •ENHANCEDST-SIZEACCURATELYANALYZES
ALLPANELCONCEPTSWITH2-DFEA
t.... rSo_aumlto_
Minimum weight
_-:_: // / /,4 ¢ ..
Figure 6. Structural Sizing Process. _"_. .::_:_,.. ./. /
In general, the structural panels of airbreathing hyper-
otmrpan_concepts •Thermoelastic
soNc vehicles are unsynunetric--geometrically
and/or thermally. As a result, traditional 2-D panel Tr_.S*r@,_ _ s_ _*__a=,=St_.._ "fFoarmiluurleatiomnosdes
methods, which do not account for panel asynmaetry, (Buckling)
can predict inaccurate panel sizes. In contrast, an _ • Composite
H°n_mb _ndwlch B_ClIst,_*_.,_ amnatderimalsetallic
enhanced version of ST-SIZE, developed by SAO, w
models the panel asymmetry. This is accomplished by Figure 7. Enhanced ST-SIZE Method.
calculating the membrane bending coupling inthe
2-D element. Thus a coarse global-sized mesh on a
complete vehicle airframe and engine, modeled with
2-D elements as shown in Figure 7, will yield the M10 THERMAL
MOMENTS
same accuracy as a 3-D subscale-sized fine mesh, __ER F(MESAC/NCAOSMTPRUATNE) D,_
even tbr unsymmetdc panels. In addition, panel con-
1% ERROR
cepts can be differentiated and selected based on their
thermoelastic formulations, failure modes and materi-
COMPUTED
als, all within a preliminary/conceptual-level design. UNIT
WEIGHTS
Figure 8compares the results of the traditional and
ST-SIZE 3.0 I_
enhanced ST-SIZE methods on the same global-sized
2-D element mesh for aMach 10vehicle. Using atradi- ;_ 12.3.o
tional 2-D panel method with MSC/NASTRAN, the TRADITIONAL METHOD psf NEW METHOD
predicted thermal moment, forexample, shows a25% Figure 8. Traditional and Enhanced ST.SIZE
error ascompared tothatfora fine 3-D subscale-sized Method Applied to a Hypersonic Vehicle _.
4
Thermal Management tanks are adjacent to the vehicle skin, fuel is
located just below Node 9. Knowing the integrat-
The following discussion on the thermal protection ed heat load into the fuel tank, the amount of fuel
system (TPS) is presented from an SSTO perspec- that must be boiled-off to maintain the tank pres-
tive, where the sizing of the TPS is dependent on sure can be determined. The transient analysis
the transient nature of the heat loading. For longer also predicts whether or not active cooling, as
flight times, such as for cruise vehicles, alternative opposed to TPS, is required for any portion of the
systems are considered. Figure 9 shows the under- vehicle surface. If at any point along the trajecto-
surface of a Mach 12vehicle color coded by the ry the temperature at Node 1exceeds the material
appropriate thermal protection thicknesses. It is temperature limit of the TPS, tor example
necessary to accurately determine the thickness of 2500°F for FRICI-12 and 2300°F for TABI, or if
the TPS to yield an accurate prediction of its the TPS thickness is greater than some prede-
weight, its volume for packaging considerations, fined maximum allowable thickness, then active
and the heat flux through the surface such that fuel cooling is required at that location on the vehicle.
boil-off rates can be determined. The heat flux into
each of the panels at several points along the tra-
jectory is known from the results of aerothermal
calculations, for example from a code like 1051 plugs (upper and lower)
11trajectory pointsconsidered for I"nrn2ukalamtl_on
S/HABR Figure 10 shows the cross-sectional view
ascent and descent _/u
950 seconds total mission _me
of one panel, or plug. Node 1is the surface of the
•Lower surfaceTPS consistsof
vehicle. The TPS is located between Node 1 and FRCI-12 tilesbonded directly to
foam tankineaJlation 0._
Node 7. Node 7 represents the bond between the •Upper surfaceTPS consists of
TABIAdvanced Blanket Insula_on
TPS and the fuel tank insulation. Below Node 9 is b_tlotdo_am_.kio_l_tlo. )ii::!L
0A
the structural panel described in the above section.
The transient analysis proceeds, as illustrated in
Active cooling
Figure 11, as follows. Knowing the heat flux at required on
Cowl leading
Node 1 from the aerothermal code at representa- edge, internal
engine surfaces,
tive points along the trajectory, and an initial
value of TPS thickness, a transient analysis is per- Figure 9. Mach 12 Staging X-34 Concept
formed starting at the initial conditions on the Thermal Protection System Thickness.
ground and marching along the trajectory. If at
some point along the trajectory, the temperature
limit at Node 7 is exceeded, for example in this
1002plugs,18ascentanddescent
case the temperature limit is set at 400°F due to
trajectorypointsanalyzed
the temperature constraints imposed by the bond-
Node 1
ing material, then the analysis is stopped, the TPS 3--
thickness is increased, and the transient analysis varmue Node 4• Nolle ;_ TUFI/FRCI-12
begins again. This process continues until the Noae _•
Node _'_ RTV 560
appropriate TPS thickness is determined such that • _ ' -175 Fto500 F
NBIBUURUJlU_Noae... _8=e_Nmomdem5m_mm_mm_o'_rTmA_BaI (Limited to 400 F
the temperature limit at Node 7 is not exceeded I1__ duets Rohacell
by the end of the trajectory. This analysis is
____ [ ____ NocTeO ___ ___ max.temp.)
repeated for each plug on the vehicle in an auto-
mated manner, where a typical vehicle is com-
ActiveCooling Increase TPSThickness
posed of over 1000 plugs.
Node 1>2500 F(FRICI-12) Node 7>400F(RTV 560)
Node 1>2300F(TABI)
Once the TPS thickness is known for each plug, TPSThk >2.0in.
the heat flux at Node 9 can be determined from Figure 10. Access-to-Space SSTO Thermal
the same transient analysis. Note that where fuel Analysis Plug Model for Sizing TPS.
Routing Schmatic L_
IDivideaircraftsurface intoplugs I
I Begintransientanalysis I<
In ,a.s.eat,oade ......[.2....!....i.11i
andcalculate/ N _-_
temperaturesthrough | Iti el IReportl_
plug.sing ! _ Pl I_ul_l-'_:_U
SINDA- 85 /
Incre_ii_nSUle_s°sn 1 A_inveeeldeCf°/inthgifi'_
• Fuel temperature/pressure to engine
•Required fuel flow for cooling
Figure 12. Cooling System Design(cid:0)Analysis.
Discipline Interdependence
Figure 11. Automated Insulation Sizing•
As previously mentioned, the areas of propulsion,
structures and thermal management are unique toair-
Generally it is known a priori that the engine
breathing hypersonic vehicle design. However the
flowpath requires active cooling. 7The upper
other disciplines are also critical to resolving a
left-hand corner of Figure 12 shows an exam-
hypersonic vehicle. Figure 13illustrates the complex
ple of a coolant routing along the keel-line of
interdependence among the disciplines in airbreath-
the inlet, combustor and nozzle. Schematically,
ing hypersonic vehicle design. For example, aerody-
the active cooling network is shown in the mid-
namics inputs surface coordinates from geometry;
dle of the figure. Inputs to the network analysis
interacts with propulsion indefining the entire vehi-
include the initial coolant system architecture,
cle configuration; outputs heat loads tothe thermal
propulsion heat loads and flowpath geometry,
management analysis; outputs forces and tempera-
coolant supply temperature, coolant and material
tures to structures; and iterates with the trajectory to
properties, and the total pressure drop through
yield flight conditions, forces and moments. As
the network, based on the pumping system and
noted previously, not only are there a large number
the desired fuel injection pressure. From this, the
of couplings, but the sensitivities are high and the
coolant mass flow, temperature and pressure dis-
system ishighly nonlinear. For these reasons, the dis-
tribution, along with the panel temperature dis-
ciplines are resolved tothe high degree of accuracy
tribution are determined. The panel temperatures described inthe sections above. This detail isneces-
are checked to ensure that they remain below the
sary just to capture the impact of the key factors in
material temperature limits. Also, panel stresses
airbreathing hypersonic vehicle design.
are calculated. For example, if a hole is punc-
tured in one of the cooling panel walls, the stress
on that wall must not be high enough to cause
the panel to "un-zip." The network architecture V_ltrne
i
and panel designs are modified until the overall
cooling system weight and coolant flow rate are
minimized, while meeting the above constraints.
As noted in the propulsion section, the coolant
flow rate and the fuel injection properties have a
significant impact on the net propulsive thrust.
choke
TPSOtt_
Figure 13. Discipline Interdependence.
6
HOLIST , flown as represented by the "Analyze Mission"
box. From the mission results, the vehicle is sized.
HOLIST isSAO's working environment for the (It isalso possible to define a scaling factor as a
multidisciplinary design, analysis and optimization variable and use IPFR-PFAI _<.1as a constraint.
of airbreathing hypersonic vehicles. It is being This would eliminate the need to perform the siz-
developed by SAO in part through a contract with ing process in the extra loop.) At this point, if only
McDonnell Douglas? HOLIST will help to elimi- a single vehicle analysis were required, the
nate disconnects between disciplines, enable rapid process would be complete. However, if it is
multidisciplinary parametrics, allow the evaluation desired to optimize the vehicle, the optimization
of design sensitivities, and will enable the optimiza- process begins. Finite differences are used to cal-
tion of the vehicle design and trajectory. Currently a culate the derivatives of the objective function
parametric geometry model, Pro/ENGINEER, is with respect to each of the design variables. Thus,
being incorporated into HOLIST. This will enable for the perturbation of each design variable, one
the entire vehicle configuration to be represented pass through the loop is made. Based on the deriv-
with anumber of specified design variables. ative information, the vehicle design for the next
HOLIST isconstructed modularly such that when iteration is defined. The objective function for the
improvements are made in any of the discipline new design is evaluated, the derivatives at the new
tools, or new tools are available, these can be easily point in the design space are determined, and the
incorporated. Auser-friendly optimizer, Optdes-X, process continues with the vehicle definition for
has been integrated into the environment. And the the next iteration. Iterations continue until the con-
entire system is set up on workstations, complete vergence criteria and all the constraints are satis-
with graphical user interfaces. fied, yielding the optimum vehicle configuration.
Figure 14 isa simplified flowchart illustrating how
an optimization proceeds in HOLIST. In the upper
left-hand comer, the process set-up includes defin- ..................._,_;_/;hfae..................i
ing the design variables, objective function, con-
I Packaging, I J Thermo I?_,,,.,,,,,,,I Optimization
straints and convergence criteria for a run. The
baseline vehicle geometry and packaging, together
with a definition of the mass and thermo properties,
follow. Analysis of the configuration proceeds with
h _ Propulsion I ...........'.1.._..,.;.;_1"1 _i'; I
aerodynamics, propulsion, etc. (Note that for sim-
plification of the diagram several disciplines are not
represented here, including structures and thermal
management, for example.) The analysis can either
be performed in real time, i.e. by running an analy-
sis code, or a database can be accessed to obtain the
discipline results. It is important to note that there is Figure 14. HOLIST Design Optimization.
more than just one result being passed through this
flowchart. In other words, since the vehicle will fly Current Status and Demonstration Example
some trajectory, matrices of aerodynamic and
propulsion data representing the coefficients of lift, Currently, the basic synthesis system of HOLIST
drag, and thrust, and fuel flow rate, for example, at is in operation. The capabilities include aerody-
appropriate values of angle of attack and Mach namics and propulsion analysis for Mach 6 to 25,
number, must be passed through the loop. In addi- and vehicle performance methods such as energy-
tion, the vehicle geometry may be variable along a state, 3-DOF and GATMIS, which can perform
trajectory requiring multiple geometry definitions. various mission segments such as cruise, maneu-
vers, descent, etc. Also included are methods for
Once the analyses are completed the vehicle is packaging, mass property definition and vehicle
sizing.Optdes-XhasbeenintegratedintoHOLIST completed trajectory and the value for the PFR.
andcanbeaccessebdyanyofthedisciplinesindi- From the mass properties and packaging, PFA is
vidually,aswellasfromthesystemasawhole. determined. Thus the objective function, PFR-PFA is
known. Using the differencing method described
Ademonstratioonftheoptimizatiocnapabilityof above the optimization proceeds with finite differ-
HOLISThasbeencompletedA.single-stage-to- ence derivatives being determined for each of the
orbitvehiclewiththebaselineconfiguration five design variables, followed by a new vehicle
shownatthetopofFigure15wasselectedA.s geometry for the next iteration. In this example the
illustratedinthelowerleft-handt,hedesignvari- number of iterations was predefmed to be twenty,
ablesincludethevehicleforebodyanglen,ozzle without the selection of a convergence criteria.
chordaal ngle,theplanformexponenatndscalar,
andtheuppersurfacemaximumheightT. hevehi-
clelengthwasheldconstanWt. iththesefivevari-
ablest,heentirevehicleconfiguratioinsdefined.
Theshapecanbeviewedon-screenc,hanging
whiletheoptimizerproceedsif,desiredT.hepri-
maryanalysersepresenteindthedemonstration
areaerodynamicpsr,opulsiona,simplifiedtrajec-
torycalculationp,ackaginagndweightsT.he
objectivefunctionwasPFR-PFAforanunsized
vehicleT. husasPFR-PFAisminimizedt,hevalue
] PFA j
will bedrivenfromapositivevalue,forexample,
towardsanegativevalue.Oncethefinalvehicleis
sizedt,heTOGWwill alsohavebeenminimized. Figure 16. Demo Problem Flowchart.
NotethatinanothearpproachT,OGWcouldbe
definedastheobjectivefunctionwithIPFR-PFAI The plot of TOGW and PFR-PFA versus iteration
_<.1asaconstraint.
in Figure 17 illustrates the results of the optimiza-
tion. The baseline configuration began with a
TOGW of 606,000 lbs with a positive PFR-PFA,
and thus an even heavier sized vehicle. After 20
iterations the final configuration had a TOGW of
i 0,o0uils,o°
389,000 lbs:with a negative PFR-PFA. Thus, if
this configui?ation is sized for the mission, the final
--a--
vehicle TOGW will actually be less than 389,000
lbs. A significant reduction in TOGW is achieved.
Variables:
1) Nozzle angle 3) Upper surface Z-constraint
2) Forebody angle 4) Planform scalar 5) Planform exponent
700 .04
Figure 15. HOLIST Demo Problem.
Baseline _..,:_:Y;".:i-i.:ii-_i:i'_>._:
_.y_::;:i::i:iiiiii_i!::'__:"
.03
Figure 16shows the actual flowchart for the demo 620 _iiiiiiiiii_ TOGW......=606K Finallterat_io.<nx+_,_2__':-?'_:_.::'-':"
problem. At the top of the figure, the geometry is .02
defined based on the five design variables. The (1000 lb.) PFR-PFA
geometry istransformed into a format that can be 46O .01
TOGW 540
read by the propulsion and aero disciplines.
380 0
Propulsion data is supplied from a database and aero- PFR-PFA
dynamic data isobtained from the S/I-IABP code 300 I5 tl0 115 20.01
while it runs in real time. The trajectory iterates with Iterations
the propulsion and aero data, finally resulting in a Figure 17. Results and Iteration History.