AD-A25 9 291 - WL-TR-92-2029 PROPULSION/WEAPON SYSTEM INTERACTION MODEL Colin Widdison The Boeing Company Boeing Defense and Space Group Military Airplane Division PO Box 3707 Seattle, WA 98124-2207 D T IC SELECTE July 1992 jA ] 5 '1993 E Final Report for Period December 1983 - September 1987 Approved for Public Release; Distribution is Unlimited AERO PROPULSION AND POWER DIRECTORATE WRIGHT LABORATORY AIR FORCE MATERIEL COMMAND WRIGHT-PATTERSON AIR FORCE BASE, OHIO 45433-6563 j. !.1 '. ~... Ij6 93-00843 K ;=' 1' I I I IJ
Transcript
AD-A25 9 291 -
WL-TR-92-2029
PROPULSION/WEAPON SYSTEM INTERACTION MODEL
Colin WiddisonThe Boeing CompanyBoeing Defense and Space GroupMilitary Airplane DivisionPO Box 3707Seattle, WA 98124-2207 D T IC
SELECTE
July 1992 jA ] 5 '1993
EFinal Report for Period December 1983 - September 1987
Approved for Public Release; Distribution is Unlimited
AERO PROPULSION AND POWER DIRECTORATEWRIGHT LABORATORYAIR FORCE MATERIEL COMMANDWRIGHT-PATTERSON AIR FORCE BASE, OHIO 45433-6563
j. !.1 '. ~... Ij6 93-00843 K;=' 1' I I I IJ
NOTICE
When Government drawings, specifications, or other data are used for any purposeother than In connection with a definitely Government-related procurement, the UnitedStates Government incurs no responsibility or any obligation whatsoever. The fact thatthe Government may have formulated or in any way supplied the said drawings,specifications, or other data, is not to be regarded by implication, or otherwise in anymanner construed, as licensing the holder, or any other person or corporation; or asconveying any rights or permission to manufacture, use, or sell any patented inventionthat may in any way be related thereto.
This report is releasable to the National Technical Information Service (NTIS). AtNTIS, it will be available to the general public, including foreign nations.
This technical report has been reviewed and is approved for publication.
AFFREYdi. STRICKER RI CHARD JC/RA L-ALProject Engineer Technical Area Manager
Aero Propulsion & Power Directorate Aero Propulsion & Power Directorate
RICHARD J. '1ILLActing Deputy for TechnologyTurbine Engine DivisionAero Propulsion & Power Directorate
If your address has changed, if you wish to be removed from our mailing list, or if theaddressee is no longer employed by your organization please notify WL/POTA, WPAFB,OH 45433-6563 to help us maintain a current mailing list.
Copies of this report should not be returned unless return is required by securityconsiderations, contractual obligations, or notice on a specific document.
f orm Approved
REPORT DOCUMENTATION PAGE OMB No. 0704-ro18PuolC, I i'o nl)n buict' tor ,hi% (Oiledtio of inlOtmal t•% M l' ti *id',.q" I hour pf resporse. in(itding the time for 0rMevwng inStNUcli o. 0e0r4hn 88q Oata sources.
gaher •i ind maint,,fli3ni the data reeded. d (0ncompleting and reviewiin th ,O t~ioiin o informatiOn Send iomment$ re arding this burden etismate or any Other asoect Of this
olletmon ot inforn-atOn. ,nluding suggestions for reduorg thou burden to W.ashington rleadquarterfs Seri.ces. Darec(soate or Informoation Operations and Repors. 12 IS Jefl•esooavis enilhway. Suite 1204. Arlington. VA 222014302. -. nd to the Office or Man"'jement and Budget. Paperwoik Reduction ProleO (0104.0188).Wasilngton, DC 20503.
1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED
July 1992 Final, Dec 83 - Sep 874. TITLE AND SUBTITLE 5. FUNDING NUMBERS
C-F33615-83-C-2347Propulsion/Weapon System Interaction Model PE - 66203F
PR - 3066
6. AUTHOR(S) TA - 11WU - 46
Colin Widdison
7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATIONBoeing Military Airplane Company REPORT NUMBER
Boeing Defense and Space GroupMilitary Airplane Division D180-29760-1PO Box 3707Seattle WA 98124-2207
9. SPONSqORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORING/MONITORING
Jeffrey M. Stricker (513/255-2121) AGENCY REPORT NUMBER
Aero Propulsion and Power Directorate (k.L/POTA) WL-TR-92-2029Wright LaboratoryWright Patterson Air Force Base OH 45433-6563
11. SUPPLEMENTARY NOTES
12a. DISTRIBUTION / AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE
Approved for Public Release; Distribution isUnlimited
13. ABSTRACT (Maximum 200 words)
This document described a computer program used to evaluate advanced airbreathingpropulsion payoffs in aircraft performance of future interest to the USAF. Theprogram is required to determine the potential impact of propulsion technologyadvancements and future weapon system requirements on propulsion concept and cycleselection. A major requirement in such assessments is the evaluation of interactioneffects between the engine and airframe. Several scalable airframe "Data bases"were developed to examine a variety of vehicle concepts including a tactical fighter,supersonic interceptor, supersonic cruise missile, logistic transport, lightweightfighter, carrier air vehicle (first stage of a two-stage-to-orbit system) and hyper-sonic interceptor. A program description including options, mission analysisapproach, installation methodology, and program structure is provided. A descriptionof each of the existing data bases including baseline weights, dimensions, anddrag characteristics is included. Also, a sample output listing is included inthis document.
hP. Enthalpy of primary exhaust flow after heat addition,Btu/lb
h Thrust height, in 2
M Mach number, dimensionless
vii
LIST OF NOMENCLATURE AND SYMBOLS (Continued)
P Static pressure, lb/in2 , perimeter, in
P, Relative pressure: the ratio of the pressures P. andPb corresponding to the temperatures T, and Tb,respectively, along a given isentrope,dimensionless
P.S. Power setting
PT Total pressure, lb/in2
Q Effective heating value of fuel, Btu/lb
q Dynamic pressure, lb/in2
R Gas constant
R,r Radius, in
RF Total pressure recovery
SFC Specific fuel consumption
SFCA Installed specific fuel consumption
T Temperature
V Velocity, ft/sec
W Mass flow, lb/sec
WBX Bleed air removed from engine, lb/sec
Wc Corrected airflow, lb/sec
Wf Weight flow rate of fuel, lb/sec
W Weight flow rate of air, primary plus secondary, lb/sec
WO Primary nozzle airflow rate, lb/sec
Tz Temperature correction factor, Tn/TsD
ST2 Pressure correction factor, PT/PS
B Burner efficiency, dimensionless
P Density, lb/ft3
vii!
SUBSCRIPTS
amb ambient
AB afterbody
b burner
B. bleed airflow extracted from the engine
BP bypass
BLC boundary layer bleed
ix
1.0 I...roduction and Summary
L.i Introduction
The Turbine Engine Division of the WL Aero Propulsion andPower Directorate frequently carries out studies to determine thepotential of propulsion technology advancements and to assess theimpact of future weapon system requirements on propulsion conceptand cycle selection. To that end, a computer program was writtento provide a rapid response capability which gives considerationto diverse mission requirements and accurate propulsion/airframeintegration.
To meet these needs the program has the followingcapabilities:
"o estimation of installation effects on engine performance
"o ability to calculate airplane performance in any user-defined mission.
The latter capability is dependent upon several items, asfollows:
"o calculation of airframe weight
"o calculation of airframe drag
"o calculation of mission performance segments such asCLIMB, CRUISE, etc
"o ability to assess the mutual influences between thedifferent technologies involved.
The program was written by adapting a set of existingpreliminary design programs into a unified program that can beused to identify airframe/mission interaction effects on advancedpropulsion systems.
Criteria for the unified program include:
"o rapid turnaround
"o interactive capability
"o simple to operate
"o modular construction.
In conjunction with the program, a data base of several"generic" aircraft configurations was provided. These genericconfigurations serve as baseline designs that can be used toassess the effects of selecting variations in engine, airframe,and installation parameters. The selection of configurationsthat have been the subject of serious study assures that theparametric analysis will be carried out realistically.
This document provides an overview of the work accomplishedduring the PWSIM contract and an introduction to the resultingcomputer programs. For detailed information, see Figure 1-1.
1.2 Summary
The objective of this research was to develop a computerprogram for the evaluation of air-breathing propulsion systemperformance in interaction with aircraft of current or futureinterest to the USAF. The program was required to allowdetermination of the potential of propulsion technologyadvancements and the impact of weapon system requirements onpropulsion concept and cycle selection. A major requirement insuch assessments is the evaluation of interaction effects betweenthe engines and airframes. The computer program was required tosynthesize a variety of vehicle concepts (Figure 1-2).
o a tactical fighter
o supersonic interceptor
o supersonic cruise missile
o logistic transport
o lightweight fighter
o carrier air vehicle (first stage of a two-stage-to-orbitsystem), and
o hypersonic interceptor.
To meet these objectives the plan of work involveddevelopment of two computer programs each consisting of anexecutive routine, two permanent modules, and an interchangeable"data base" module. Two programs were necessary because of theunique mission requirements for the carrier air vehicleconfiguration and the desiqn implications imposed on it by themission requirement of the second-stage vehicle.
w
(3 0 Q CAILL F- z! OC S
- . o- 5~ 0 W-cr
I- W Cf) w .
-wa:- w- Ln4
> -)
U) -
C')~ cc0 j~
0 >~c0 g:D U)
C* CO0)c z:<<.ODw > I
E,~F--
_ý 0
(1)in )
cc Ca i0 W > tw0-2o
a. yCO) n cc
In
I.I.-
UL
CL 0
I Lhus - H
- 4.
U,
00
600
0r
-.. amc
0
cc
W.. I
LU~
These programs were mainly derived from several already inuse at Boeing, and the bulk of work was associated with adaptionand integration of these programs into distinct, compatiblemodules.
The two permanent modules calculate engine installationeffects and airplane size and performance, respectively. To meetthe objective of realism (in a preliminary design sense) sevendata base modules were developed for use with the program; theserepresent seven "generic" aircraft configurations. Each database module contains a data description of a baselineconfiguration and several routines that allow the program user toscale and modify the baseline with an input data file.
To evaluate the potential of propulsion technology advances,it is necessary to measure their effect on the performance of thesystem they are likely used in. To be useful, such assessmentsmust be made at the very early stages of technology developmentto identify promising approaches. Thus, there is need for arapid response capability which gives consideration to a widerange of configurations, diverse mission requirements, andaccurate propulsion/airframe interaction assessment.
The computer program that was developed is capable ofcalculating either the mission performance of an aircraft ofknown size or the size of such an aircraft required to perform aspecified mission. In executing these calculations, the programtakes account of the engine performance (as supplied by theengine manufacturer), engine installation effects of the inletand nozzle on the engine performance, other engine/airframeinteractions, body volume dictated by engine dimensions, wingsize influenced by fuel tank volume, etc. and missionrequirements. The two main modules of the program deal withengine installation analysis and airplane performance,respectively.
The propulsion installation module is a simplified versionof the Boeing Engine Installation Analysis Program (EIAP) andcalculates the inlet, nozzle, and aftbody effects on theuninstalled engine performance data supplied by the enginemanufacturer.
Inlet effects considered are inlet drag, inlet recovery, andthe effects of mismatched inlet air supply and engine air demand.In addition to the effects on engine performance, the inletroutines also allow evaluation of the inlet capture area; thisinformation is supplied to the d-qta base routine to allow enginedimensional considerations to be accounted for in the evaluationof airframe geometry, drag, and weight. Exhaust considerationsincluded are gross thrust and aftbody drag effects.
6
The airplane performance module (incorporating an airplanesizing option) allows the evaluation of mission performance foran aircraft of known properties (including engines of knowninstalled performance) and also allows "point performance"evaluation at user selected values of weight, altitude, and Machnumber.
The mission analysis is performed by calculating aircraftperformance during distinct segments (CLIMB, CRUISE, COMBAT,etc.) and linking these segments into a complete missiondescription.
Seven data base modules were developed for use with thecomputer program. Each data base represented a typicalconfiguration and was based on an actual preliminary designstudied at Boeing.
The data base module consists of:
1. Baseline configuration and modification module thatdefines the baseline configuration and allows the userto modify (with input data) many of the aircraft designparameters (wing loading, aspect ratio, etc.)
2. A geometry module that evaluates the aircraftdimensions, inlet size, fuel volume, etc.
3. A drag evaluation module that constructs drag tables foruse in the performance module, and
4. A weight module that calculates the fuel and operatingweights of the aircraft of known gross weight andpayload and feeds these numbers to the performancemodule.
2.0 Program Description
The Turbine Engine Division of the WL Aero Propulsion andPower Directorate is continually engaged in internal studies todetermine the potential of propulsion technology advancements andto assess the impact of future weapon system requirements onpropulsion concept and cycle selection. This information isrequired to support Division Long-Range Propulsion planning andprogram resource allocation in the exploratory and advanceddevelopment areas.
To fulfill the need for a rapid response capability whichgives consideration to diverse mission requirements and accuratepropulsion/airframe integration (and thus provide timely
7
technical assessment program planning information), acomprehensive automated evaluation process is required.
Large, complex computer programs have been developed by theaircraft and propulsion industries to assess future systemrequirements on advanced weapon system designs. Because of theirlarge computer storage requirements, extensive data base needsand long execution times, these programs could not be efficientlyused by the Aero Propulsion and Power Directorate for advancedpropulsion assessment work.
The program described in this report combines features ofseveral existing programs for preliminary conceptual analysisinto a single program, tailored to specific needs for in-housepropulsion assessment. The program is small enough for useinteractively but retains sufficient detail in the engineeringcalculations it performs to assess the effects of engineinstallation, airframe size and geometry, and missionrequirements.
PWSIM is an interactive program for assessing the effects ofdifferent engine cycles, engine installations, missionrequirements, and airplane geometry on airplane size and weight.
The program is presently able to support seven genericaircraft types, but due to its modular construction, it canaccommodate additional configurations.
Configurations currently supported are:
o Tactical Fighter
o Supersonic Interceptor
o Supersonic Cruise Missile
o Long Range Transport
o Lightweight Fighter
O Carrier Air Vehiclt
o Hypersonic Interceptor
The program has been coded in extended FORTRAN 77 and runson the CDC Cyber 175 computer under the NOS 2 operating systemwith a required field length of about 220K octal words.
8
The complete program is stored in several differentpermanent files: one containing the main program executive andthe others consisting of libraries of modules which are accessedby the executive. The executive routine accepts the user's inputdata and controls the sequence of operation to obtain engineperformance data and then evaluates airplane size and missionperformance. The library modules are of three types:
o propulsion library
o performance library
o data base library.
The propulsion library file contains the routines requiredto read the uninstalled engine performance data and the inlet andnozzle characteristics and then performs the necessarycalculations to evaluate the installed engine performance. Theperformance library contains the modules needed to calculate thepoint performance and mission performance of an airplane derivedfrom a data base library.
Several data base libraries are available. Each data baselibrary contains all of the configuration related modulesrequired to define and scale the geometry of a baselineconfiguration and evaluate its drag polars and operating weight.To accommodate a new configuration, it is necessary to create anew data base library. To execute the program, it is necessaryto LOAD the executive program with the performance and propulsionlibraries and with an appropriate data base library. Also neededto execute the program are several sets of input data; most ofthe input data are stored in permanent files, but the user has acertain amount of control via interactive inputs. When theprogram is executed interactively, prompting messages areprovided to the user at the terminal. These messages serve as aguide to allow proper selection of the input required. Most ofthe interactive inputs are for either selecting calculationoptions or identifying input data files; some interactivenumerical inputs are required when the engine installation optionis selected.
2.1 Basic Options
The principal features of the program are:
o An engine installation module that converts enginemanufacturer's uninstalled engine performance data intoinstalled performance data by evaluating the internallosses and drag characteristics for the inlet andnozzle/aftbody configuration.
9
o A set of data files containing inlet and nozzle/aftbodyperformance maps applicable to suitable engineinstallation configurations.
o A set of data modules containing data definitions of thegeneric airplane configurations which allow assessmentof the geometric, aerodynamic, and weightcharacteristics of scaled versions of a baselineaircraft.
o Technology modules that provide rapid and reliableestimates of airplane drag and airframe weight.
o A mission analysis module that allows the user to definealmost any practical mission.
o Mission segment modules that use the installed engineperformance data and calculated drag polars togetherwith accepted performance methods to assess time, fuel,and distance required to complete the segment.
o The option to scale or "size" an aircraft for a givenfixed mission or to find the cruise range or loiterendurance of an aircraft of prescribed size.
o The option to use the engine installation module or useof previously installed engine performance data whereappropriate to reduce computation time and cost.
o A choice of interactive or batch operation.
o An output file providing graphics data for aconfiguration drawing.
This section describes the capabilities and overalloperation of the program and outlines the options available tothe user.
The program's principal function is to calculate airplanemission performance for an airplane derived from one of severalbaseline designs.
2.1.1 Mission Performance
Several important parameters affect the performance of anaircraft, namely:
dragweightand propulsion system performance.
l0
For an aircraft of known size and geometric proportions, thedrag and weight can be estimated (to a satisfactory degree ofaccuracy) using methods combining analytical and empiricalrelationships among certain important design parameters.
With such a design the propulsion engineer has an extremelyuseful tool for assessing the payoff obtained from gains inengine performance, comparing the performance levels of differentengines or evaluating the impact of propulsion concept and cycleselection.
The performance level of the propulsion system is not solelya function of the engine or engines selected; it is also stronglydependent upon the way in which the engines are combined with theairframe. For this reason, it is important to assess the effectsof the inlet and exhaust systems on the net thrust produced bythe engine at any flight condition of interest.
The Propulsion/Weapon System Interaction model computerprogram (PWSIM) has the capability of evaluating drag, weight(and, therefore, fuel load) and propulsion system performance(including installation effects). It has the furthercapabiiities of calculating airplane performance in terms of thebasic components of mission performance such as climb, cruise,loiter, etc. (referred to in this document as mission segments).
In addition to the above, the program provides a facilityfor calculating the performance of a mission composed of a stringof mission segments selected by the user.
Two modes of mission performance are available to the user:
"o the aircraft begins the mission at a specified weight(and fuel load) and the program calculates the extent ofthe mission, the end being determined by attaining aweight equal to the sum of the operating weight, anyremaining payload and a specified amount of reservefuel. (Note: the extent of the mission can haveseveral meanings since the mission requirements mayinclude varying amounts of cruise or loiter.)
"o the requirements to achieve a given fixed mission maydictate the use of more fuel than can be carried by anassumed baseline design. The so called "SIZING" optionof the program allows scaling of the baseline aircraftto accommodate the extra fuel while taking into accountthe resulting inrrea~s in weight, drag, and enginesize.
II
Figure 2.1-1 shows the interrelationships among thedifferent technologies involved in assessing airplaneperformance; the message in the box in the lower right cornerindicates the two modes of matching the airplane to requirements.
2.1.2 Baseline Designs
The items in Figure 2.1-1 that occupy rectangular boxes areall configuration-dependent. In order for the program to supporta wide variety of configurations, each of these areas of thecomputer program would require sufficient input data (definingthe configuration geometry and its influence on drag and weight)to differentiate among the various techniques and methodologiesrequired for the different configurations. Program modulesdesigned to accommodate a wide range of configurations would havethe further disadvantage of being complicated because of thelarge number of decisions required in selecting an appropriatesequence of calculations which results in long execution timesand difficulty of maintenance.
In PWSIM these disadvantages have been minimized by the useof "baseline" configurations that have been coded into theconfiguration-dependent parts of the performance analysisprocess. By this technique the amount of input data required todefine the airframe geometry, weight, and drag is minimized andcorresponding program logic is kept relatively simple. Asufficiently large input data set is retained to allowconsiderable variations from the baseline configuration both ingeometry and application.
The apparent disadvantage of being limited to specificconfigurations is easily overcome because the technologydependent program logic is contained in modules that evaluateairplane geometry, drag and weight, respectively. Thus, analternate baseline can be "swapped-in" to the program withrelative ease.
The program currently supports seven baseline designs:
"o Tactical Fighter
"o Supersonic Interceptor
"o Supersonic Intercontinental Cruise Missile
"o Long Range Logistic Transport
"o Lightweight Fighter
12
U))wC
0,()
(0(j
ww LL Z l
zz
0w
L z 13
o Carrier Air Vehicle
o Hypersonic Interceptor
Each was designed by a team of experienced preliminarydesign engineers to meet specific requirements and as suchrepresents a reliable "point-of-departure" for parametric studiesof engine/airframe interactions.
2.2 Mission Analysis
This section describes the method of defining missionprofiles.
2.2.1 Background
In many simple performance programs, a mission profile isset up by coding a special subroutine to handle the scheduling ofmission segment subprograms, to pass each segment of the inputdata it needs, and to receive from each segment the output valuescomputed. Complete flexibility as to number, type, and sequenceof segments can be achieved in this way; however, the process ofsetting up or changing a mission program coded in this way isquite cumbersome, since every change requires recompiling andredebugging. In addition, such programs rapidly become expert-dependent, due to the extensive prior knowledge required of theprogrammer.
Instead of being coded into separate subroutines, missionprofiles are defined by a set of input records. At executiontime subroutine MISSION schedules segment calculation in theproper sequence, transfers data between segments and handles anyiterations required to compute mission distance, time or fuel.All these function are made transparent to the user.
2.2.2 Missions
WHAT A MISSION IS
In the current context, a mission is a flight path thatdescribes the intended usage of the airplane. Missions can beseparated into two main classes: fixed performance and variableperformance. In a fixed performance mission, all distances andtimes are fixed, and the result to be computed is the requiredfuel. In a variable performance mission, the available fuel isfixed and either the mission distance or the mission endurance isto be computed. Variable performance missions are furthersubdivided into three categories:
14
"o Range Missions - The airplane takes off and lands atdifferent places. The computed range is the distancebetween the takeoff and landing points.
"o Radius Missions - The airplane takes off and lands atthe same place. The computed radius is the maximumseparation distance of the airplane from the takeoff andlanding point.
"o Endurance Missions - All distances in the mission arefixed. The result to be computed is the time that canbe spent aloft.
PWSIM has the capability of computing all these types ofmissions.
HOW A MISSION IS DEFINED
For analysis, the mission profile is broken down into anumber of distinct maneuvers called mission segments.Performance within each segment, that is the distance, time, andfuel required to perform the desired maneuver, is computed fromappropriate simplification to the full equations of motion.Individual segments are linked end to end to approximate thedesired flight path.
Consider the sample mission profile illustrated in Figure2.2-1. This profile is typical of high-low-low-high rangemissions, with cruises of various lengths performed first ataltitude then at sea level, and finally at altitude again.Climbs and accelerations are performed for gaining altitude andspeed, but no distance credit is taken for descending ordecelerating. A total of nine segments are used to describe thisflight path.
The data needed to define a mission segment is illustratedin Figure 2.2-1. In general, the definition includes the segmenttype, the available thrust, the initial and final operatingconditions (Mach number and altitude) and for segments such ascruise and loiter, the length of the segment. Other segments,such as acceleration and climb, have lengths determined by theinitial and final Mach numbers and altitudes. The performance ofthe airplane in any one seqment is a function of the segmentdefinition and the airplane weight in that segment and hence ofthe position of that segment in the mission profile sequence.
The mission shown in the example is a variable performancerange mission i.e., the amount of fuel in the airplane (at thestart of the mission) is fixed and the total mission distance is
15
7/
1. Takeoff allowance: 5 minutes at intermediate power.
2. Accelerate from Mach 0.3 to Mach 0.5 at sea level.
3. Climb to 25,000 ft. and Mach 0.75.
4. Cruise at 25,000 ft. and Mach 0.75. Elapsed distance at theend of this segment is 40% of the total range.
5. Cruise at sea level and Mach 0.80 for 75 nmi.
6. Cruise at sea level and Mach 0.70. Distance is to be fixedby the total range capability.
7. Climb to 30,000 ft. and Mach 0.70.
8. Cruise at 30,000 ft. Total distance for segments 7 and 8 is200 nmi.
9. Loiter 20 minutes at sea level and Mach 0.35.
Figure 2.2-1. Sample Mission Profile
16
to be computed. The distance covered in all segments will befixed by the segment definition except for segments 4 and 6.These cruises are free to expand and contract until the requiredmission fuel agrees with the fuel available in the airplane.Distance will be divided up between these two segments so that40% of the total range is covered before the end of segment 4,and 60% is covered after.
The appearance of a variable length cruise is indicative ofa variable performance range or radius mission. A radius missionmust have at least two variable distance cruises, of course,since both the outbound and return leg distances are to becomputed. For a variable performance endurance mission, allcruise distances must be fixed, and exactly one loiter segmentmust have a variable time. The extent of the loiter will then becomputed so that all available fuel is consumed. For a fixedperformance mission, all segments have a fixed length.
2.2.3 Mission Segments
A mission profile is defined to the program by setting up amission definition file. This file consists of a number ofrecords, each one of which defines a single mission segment. Thesequence of the mission segment records in the mission definitionfile determines the sequence in which the mission segments areexecuted to form the mission profile. The airplane weight at thestart of one segment is set equal to the airplane weight at theend of the preceding segment.
The program has a library of modules to compute fuel used indifferent types of mission segments including:
o TAXI operating at a fixed Mach number, altitude, andpower setting for a fixed period of time.
o TAKEOFF accelerating from a standstill to a prescribedpercentage over stall speed and climbing to aprescribed height.
o ACCEL accelerating at constant altitude and powersetting from the initial to the final Machnumber. Positive or negative acceleration isacceptable.
o CLIMB climbing from the initial to the final altitude.Available climb schedules include constantequivalent airspeed, constant Mach number, or acombination of the two. Climb schedule may beselected by the module for best rate of climb.
17
"o CRUISE performance may be computed for constantaltitude cruise or Breguet-type climbing cruise;cruise Mach number and altitude may be selectedby the user or computed for best range factor.
"o REFUEL transfer fuel for tanker to primary mission A/P.Tanker performance simulates the KC-135A.
"o COMBAT performing a prescribed number of max sustainedg-turns. Maneuver load factor may be set bystructural limits, maximum lift coefficient oravailable thrust and may be altered by transferfrom the initial to the final Mach number andaltitude.
"o DESCENT dropping from the initial Mach number andaltitude to the final Mach number and altitude.
"o LOITER loiter may be performed at constant altitude orconstant lift coefficient. Mach number andaltitude may be specified by the user or may becomputed for optimum endurance factor.
"o DROP dropping payload, fuel tanks, or otherwiseintroducing a weight discontinuity to themission profile. Drag for the items dropped maybe changed by selecting an index that selectsone of five arrays of additional drag vs. Machnumber.
These are the segments that may be linked together to approximate
the mission profile.
HOW A MISSION SEGMENT IS DEFINED
In the most general case, the following data are required tofully define a mission segment.
"o segment type defines the basic rules governingperformance calculation. Examples are:TAXI, TAKEOFF, ACCEL, etc.
"o power setting refers to a thrust index number definedin the engine deck. For some segmenttypes this index number defines theactual thrust used: TAXI, TAKEOFF,ACCEL, CLIMB, COMBAT and DESCENT. Forother segment types, CRUISE, REFUEL andLOITER, this index defines the maxavailable thrust.
18
"o extent defines the duration of some segmenttypes. May specify time for LOITER, fueltransferred for REFUEL, time for TAXI,distance for CRUISE, or number ofcomplete turns for COMBAT. For othersegment types (TAKEOFF, ACCEL, CLIMBDESCENT) the segment duration is governedby the initial and final Mach numbers andaltitude.
"o initial Mach number
"o initial altitude
"o final Mach number
"o final altitude
WHAT THE MISSION SEGMENTS DO
The following paragraphs describe the function of themission segment performance modules.
In several of the SEGMENTS described below, a flag (TLIMIT)is set to 1.0 if there is insufficient thrust available toachieve the required performance. When program control returnsfrom the SEGMENT calculation to the MISSION subroutine, thisvalue of the flag (TLIMIT) causes printout of the mission historyto halt at this segment and print a message to that effect.
TAXI
The TAXI module computes the amount of fuel required tooperate at the specified Mach number, altitude and power settingfor the time specified (in hours). In this and all subsequentsegments, zero is not a valid Mach number.
TAKEOFF
The TAKEOFF segment approximates the time and fuel used intakeoff, that is, between brake release and the end of climbout.Takeoff is approximated by a two-part acceleration to 120% ofstall speed, where stall speed is determined by the configurationdefinition variable CLMAX.
The first part of this acceleration approximates the groundroll up the lift-off speed which is 110% of stall speed. Anaverage acceleration is computed at 0.707 of lift-off speed usingthrust determined by the specified power setting and drag
19
computed at lift coefficient CLG, a configuration definitionvariable. A drag increment for landing gear, Figure 2.2-2, isincluded.
The second part of the takeoff acceleration, from 110% ofstall speed to 120% of stall speed, approximates the climbout.The average acceleration is computed at 115% stall speed, atspecified power setting and 1-g lift coefficient. The gear dragincrement is included over half this segment.
ACCEL
The ACCEL segment computes the distance, time and fuelconsumed in a constant altitude acceleration (or deceleration)between the specified initial and final Mach numbers. Thrust iscomputed at the input power setting.
Values for altitude and initial Mach number must be suppliedto segment ACCEL; however, this segment has two options forcomputing final Mach number. If "MAX" is specified in place ofthe final Mach number the segment computes the termination Machnumber from placard-limit or thrust-limit conditions. If "MIN"is specified, the final Mach number is computed from the stallmargin or thrust limit. Configuration definition variablesrequired to use these options include:
ZMSLM Max sea level Mach number; defines the constantequivalent airspeed part of the placard.
ZMSUP Max Mach number at altitude; defines the constantMach number part of the placard.
CLMAX Takeoff configuration max lift coefficient.
CLMAXF Ratio of landing configuration max lift coefficientto takeoff configuration max lift coefficient.
An exceptional condition occurs when insufficient thrust issupplied for acceleration or excess thrust is supplied fordeceleration. This exceptional condition is flagged by the ACCELsegment by setting flag TLIMIT=1.
CLIMB
The CLIMB segment computes the distance, time and fuel toclimb from the initial Mach number and altitude to the final Machnumber and altitude. Thrust is computed at the specified powersetting.
The climb schedule used is determined by the specified Machnumbers and altitudes, as shown in Figure 2.2-3. First, if theequivalent airspeed at the specified final Mach number andaltitude is greater than the equivalent airspeed at the specifiedinitial conditions, Figure 2.2-3a, then the climb is performedholding equivalent airspeed constant at the initial value. Inthis case, the computed final Mach number may not be equal to thespecified final Mach number. Second, if the specified finalequivalent airspeed is less then the specified initial equivalentairspeed and if the specified final Mach number is greater thanthe specified initial Mach number, Figure 2.2-3b, then a two-segment climb is performed; a constant equivalent airspeed climbis performed until the Mach number equals the final Mach numberand then a constant Mach number climb is performed until thealtitude is equal to the final altitude. Finally, if thespecified final Mach number is less than the specified initialMach number, Figure 2.2.3c, then the climb is performed holdingMach number constant at the specified initial value. Here again,the calculated final Mach number may not agree with the specifiedfinal Mach number.
The climb schedule may be optimized by specifying "OPT" inplace of either the initial or final Mach number (or both). Thequantity for which "OPT" was specified will then be computed soas to maximize rate of climb at the specified altitude and powersetting. Climb schedule determination will then proceed asabove.
An exceptional condition occurs when the airplane becomesthrust-limited along the climb schedule before reaching the finalaltitude. This condition is flagged by setting the flagTLIMIT=I.
CRUISE
The CRUISE segment computes the time and fuel required tocruise the specified distance. The specified power settingdefines the maximum available cruise thrust.
The cruise Mach number may be a specified constant or may beoptimized by the segment. If a constant is specified, Machnumber is held fixed at this value throughout the cruise. If"OPT" is specified in place of the initial Mach number, both theinitial and final Mach numbers are optimized independently, and aslight acceleration or deceleration might result.
Three options are available for determining the altitudeprofile of the cruise. If the initial altitude is a specifiedconstant and the final altitude is not the same constant, then a
Breguet-type climbing cruise is performed. Next, if the initialand final altitudes are specified equal to each other, than aconstant altitude cruise is performed. Finally, if "OPT" isspecified in place of the initial altitude, then both initial andfinal altitudes are independently computed so as to maximizerange factor.
If insufficient thrust is available for cruise, the segmentvalue of the flag TLIMIT is set to 1.0.
COMBAT
The COMBAT segment computes the time and fuel required toperform the specified number of 360-degree turns at max sustainedload factor. Thrust is computed at the specified power setting;load factor is computed so that the resulting drag agrees withavailable thrust. The required drag may be increased (ordecreased) by the specific energy released (or consumed) intransferring from the initial to the final operating conditions.If final conditions are not specified, they are taken to be thesame as the initial conditions.
If insufficient thrust is available at the specified power
setting, this is indicated by the flat TLIMIT=1.0.
DESCENT
The DESCENT segment computes the distance, time, and fuelused in descending from the initial Mach number and altitude tothe final Mach number and altitude. During the descent, the rateof change of speed with altitude is held constant.
Optionally, the user may specify "MIN" in place of the finalMach number. In this case, the final Mach number will either be120% of the landing stall Mach number (as determined byconfiguration definition variables CLMAX and CLMAXF) or the lowerthrust limit Mach number, whichever gives the higher value.
REFUEL
The REFUEL segment computes the distance, time, and fuelconsumed in receiving the specified weight of fuel from a tanker.The fuel transfer rate and downwash velocity simulate the KC-135Atanker; however, no check is made as to whether the KC-135A couldoperate at the specified Mach number and altitude or couldtransfer the required weight of fuel.
The user may specify "MAX" in place of the weight of fuel tobe transferred. This signals the segment that fuel is to be
24
transferred until the weight of the airplane is brought up toTOGW, a configuration definition variable.
The final segment weight is restricted to the weight atwhich the airplane would become thrust limited at the segmentMach number, altitude and power setting. The segment limits thefuel transfer to no more than the amount that would bring theairplane weight up to the thrust limit weight.
LOITER
The LOITER segment computes the fuel required to operate forthe specified number of hours at the operating conditionsdefined. The specified power setting defines the maximum thrustavailable.
The loiter Mach number may be a specified constant or may beoptimized by the segment. If a constant is specified, the entireloiter is performed at this Mach number. If "OPT" is specifiedin place of the initial Mach number, the initial and final Machnumbers are optimized independently and a slight acceleration ordeceleration may result.
Three options are available for determining the altitudeprofile of the loiter. If the initial altitude is given as aconstant and the final altitude is not given as that sameconstant, then the segment is performed holding W/8 constant atthe initial value. Second, if the initial and final altitudesare the same constant, then the entire segment is performed atthe constant altitude. Finally, if "OPT" is specified in placeof the initial altitude, then both the initial and finalaltitudes are optimized.
If insufficient thrust is available to perform the loiter,then the segment thrust limit flag is set (TLIMIT=i.).
DROP
The DROP segment provides a way to introduce a weight ordrag discontinuity into the mission profile. When the DROP isencountered, weight is decremented by the specified number ofpounds and drag is incremented by a value found in one of fivearrays of D/q vs. Mach No. selected by the value of INDEXST. Nodistance, time, or fuel is used by a DROP segment.
2.3 Propulsion Installation Methodology
2.3.1 Introduction
The Engine Installation Analysis Program (EIAP) has beendesigned to execute on the Boeing Computer Services (BCS) EKS
25
computer system. It is an overlaid program, which is writtenentirely in FORTRAN IV, and occupies 130K core locations whenresident in the computer. The program is interactive in naturein that it asks the user questions in order to input data and toattach files needed for execution.
The program is a suboverlay of the Propulsion/Weapon SystemInteraction Module, that computes installed engine performancebased on a set of engine library maps and an "uninstalled" enginedata file. The map library consists of sets of inlet and nozzleperformance data. Section II contains an overall description ofthe installation program and a discussion of the procedures usedto calculate inlet performance, nozzle performance, and theinstalled gross thrust. This manual also contains a macro flowchart of the installation module and detailed description of theprogram subroutines.
The execution of EIAP is discussed at length in the EIAPUser's Manual. The manual describes the program's interactiveinputs that are required from the user, as well as the tables ofinlet and nozzle performance, uninstalled engine data, and dragreference conditions, which must exist prior to execution. Ingeneral, the interactive inputs are used to select the following:
1. file of uninstalled engine data to be processed
2. inlet performance maps from map data base
3. nozzle performance maps from map data base
4. inlet capture area sizing criteria
5. nozzle type (axi or 2-D, convergent or con-di) and limiton nozzle exit area (optional)
6. file of drag reference conditions
7. output options.
2.3.2 Structure and Usage
The engine installation analysis program was designed tospeed up the process of calculating installed propulsion systemperformance data while including realistic effects of inlet andnozzle losses due to drag and internal performance. The programwas also designed to satisfy two additional criteria: (1) theaccuracy of the data generated by the calculation procedure mustbe suitable for use in preliminary design studies (when detailedknowledge of all geometric features of the design are not known)
26
and (2) the method must reflect the effects of throttle-sensitivechanges in inlet and nozzle/aftbody losses.
EIAP was developed from previous propulsion systeminstallation programs. EIAP utilizes a computer-stored libraryof inlet and nozzle performance characteristics and uninstalledengine data as input to interactively calculate installedpropulsion system performance. A chart showing how this computerprogram is used in a typical preliminary design analysis processis presented in Figure 2.3-1. The calculation of installedpropulsion system performance is almost instantaneous if thetabulated performance characteristics of the desired inlet andnozzle/aftbody configurations are available are previously-storedcomputer files. To provide a readily-available source of inletand nozzle/aftbody data, a library of inlet and nozzle/aftbodyperformance characteristics was created that covered a widevariety of possible configurations. During execution of EIAP,these files are attached externally to the program. The userthen enters the interactive input commands. The output from theprogram can be displayed on a terminal or stored on a output diskfile for disposition to an off-line printer.
The single most important factor that made it possible toreduce the time required to perform installed propulsion systemperformance calculations was the extensive use of computerizedfiles. These files contain tables of data representing thenondimensionalized performance characteristics of inlets andnozzles. They allow instant retrieval of inlet andnozzle/aftbody data that can be matched with the uninstalledengine performance data (also contained in a computer file)during the execution of the program. The formats of the inletand nozzle/aftbody computerized files and the uninstalled enginedata were selected to provide a standardized frame work in whicheither experimental data or the results of analyticalcalculations could be used. The input format for the dataremains constant, but the data that go into the tables can comefrom various sources depending on the amount of time availablefor preparing the data and/or the amount of experimental dataavailable. Because data in the input tables can be changed asbetter data become available, it is possible to improve theaccuracy of the installed propulsion system performancecalculations as the aircraft development cycle progresses frompreliminary design through full-scale flight test.
The installation module consists of calculations that fallinto one of two main categories. The first, the inlet procedure,handles the functions of sizinq the inlet, matching the inletinput data with engine airflow demand, and obtaining the matchedinlet performance parameters from the inlet data tables. Engine
27
z.=) U
OLn
-J
< m5z cc
0 w
0 cc
CLLcc w
I- 0
0u cc
cr7 a..(LW
z
00LLD _
CLL
Coz
:)a:
z L
C:ui4I/28D
corrected airflow is the matching parameter between engine dataand inlet data. The second category, the nozzle procedure,handles the calculation of the nozzle/aftbody drag and nozzleinternal performance. Nozzle pressure ratio, PTS/Po, is used asthe matching parameter.
INLET PROCEDURES
The inlet performance procedure of EIAP is considerablylonger than the nozzle performance procedure. This is becausethe individual inlet component drags that contribute to the totalinlet drag must be calculated separately. Each of these drags(spillage, bleed, and bypass) must be determined individually asa function of mass flow ratio, which adds to the complexity ofthe computer program.
INLET PERFORMANCE
The inlet performance maps are input to the program prior tothe call to the inlet procedure. This procedure sizes the inletcapture area (if it is required) and converts the inletperformance maps into total pressure recovery and inlet dragsthat are matched to the corrected airflow demands of the engine.
The operation of the inlet procedure is shown schematicallyin Figure 2.3-2. The connecting link between the engine data andthe inlet procedure is engine operation at a desired inlet massflow ratio and recovery using the design engine airflow demand.A specified capture area size can be input, if desired, insteadof requiring the program to calculate the size.
The inlet input requires three tables of input data whichdescribe the performance characteristics of the inlet.Engineering data obtained from wind tunnel tests and theoreticalcalculations are used to obtain the inlet performancecharacteristics. The format of the inlet tables is shown inFigure 2.3-3. The nomenclature for the tables is shown in Figure2.3-4. Together, the tables form a map, which is entered intothe EIAP map library.
The inlet procedure recognizes two modes of inlet operation:low speed mode and high speed mode. The low-speed mode is usedonly at very low Mach numbers, e.g., takeoff conditions, whenonly high engine power settings are likely to be of interest andinlet drag is negligible. The high speed mode is used over theremaining Mach number regime. The EIAP calculations of recoveryand drag are illustrated in Figure 2.3-5. The requiredperformance maps are input as tables, as indicated. In this modeand the low-speed mode, recovery is read directly out of Table
29
f ~Mo DES ,Wc ,ENGDES CSEC DES
r - - -Mo ALT, THROTTLE SETTINGI f
INLET EINLETREPRESENTATION SIZING• ENGINE
(INPUT DATA SUBROUTINE PROGRAMTABLES) PRGA
0WCENG,
WCSEC,
IPTAIRFLOW 0
MATCHING
SUBROUTINE
A A01 0 B
XA A'
I ~C C
A PT0Bp 2IA pe TO0
.. • INLET
PERFORMANCE
SUBROUTINE
PT
22 INLET
T 0
Figure 2.3-2. Inlet Procedure
30
Table 103 Table 104filbymp Inlet MsoiduRuoom" Mm Flow
103 as a function of local Mach number only. The inlet drag,including spillage, bleed, and bypass drag, is found in Table 140as a function of Mach number and the ratio of inlet correctedairflow and inlet capture area (WCAC). WCAC is calculated fromthe engine demand, inlet recovery, and the inlet supply mass flowratio (found in Table 104 as a function of Mach number). Theminimum Mach number entered in Table 140 is used as the minimumvalue for which the high speed mode is used.
If the corrected airflow delivered by the inlet isinadequate to meet the engine demand at the scheduled recovery,the program will permit the inlet to operate at an excessivesupercritical margin. The recovery will be lowered sufficientlyto match the engine corrected airflow demand, and an appropriatemessage will warn the user of an undersized inlet.
INLET SIZING
The inlet sizing procedure in the computer programdetermines the inlet capture area required to match the largestengine airflow demand at each Mach number. From these calculatedinlet sizes, the largest required size is selected as the inletcapture area. For sizing calculations, an input curve (Table104) of recommended (matched) inlet airflow variations (Ao/A.)vs. M, and an input curve (Table 103) (see Figure 2.3-3 for thesetables) of recommended (matched) inlet total pressure recoveryvs. M, are used to determine the required capture area variationwith Mach number. These parameters are used in the followingequation to calculate area, A,:
WV' 0 P___ A
AC' in2 = AoENG Pr OrT A*0
(A,/A,) MATCHED 0. 343 (AO/Ac) MATcHED
INLET RECOVERY CORRECTION
The engine input provides the required data for inlet drag,inlet recovery, nozzle/aftbody drag, and nozzle coefficientcalculations. The engine section of EIAP calculates only thechanges in internal performance due to changes in inlet recovery.Changes in inlet recovery produce a directly proportional changein nozzle pressure ratio, airflow, and fuel flow because thenozzle throat area does not change. Furthermore, it is assumedthat engine data are calculated with MIL-STD-5008B recovery. Allinlet recovery changes are made relative to that value unless theuser inputs a different reference recovery. Thermodynamic datafrom Keenan and Kaye tables have been "curve-fitted", andsubroutines are provided to calculate the thermodynamicproperties of the exhaust gases.
33
The gross thrust calculation procedure is as follows: foreach altitude, Mach number, and power setting, the net thrust(FN), fuel flow (W.), corrected airflow (W40 2/6 2), nozzle throatarea (A.) , nozzle exit area (A9), and nozzle thrust coefficient(CFG) are given.
A standard atmosphere and MIL Standard 5008B inlet recoveryare used to calculate the airflow at the engine face. Grossthrust is found for the given engine data (before any changes ininlet recovery) by the following equations.
W2vF GoL, = FN + Wg
The desired inlet recovery is obtained from the inletprocedure, and the engine gross thrust is first calculated withMIL Standard recovery and then with the calculated recovery. Tocalculate engine gross thrust, the engine corrected airflowremains constant for any change in inlet recovery, and at anygiven power setting, the nozzle exhaust areas and burner fuel-airratio also remain constant. The engine performance for anychange in inlet recovery is calculated by the followingrelations:
= (PT,/ PTo)(W8) RF =W8 (P T./PTO)
(PT/PTO MIL 5008B
(STT/ PTo )
(WF) =WF (PT?/PTO)
(P _/ T.) MIL 5008B
(W)F= W2 (P.%/P•)
SHTL 500AB
(P T. / PT)
MT h 5008B
(RF - Recovery factor)
After the above quantities are computed, the correctedquantities (We) R,, (Wr) 3, (W2 )O and (PTe/Po) , are used to compute anew gross thrust, F02. This new gross thrust and the grossthrust, F. 1,calculated using the same subroutines and theuncorrected (MIL 5008B) quantities (WA, W , W2, PTO/Pj) are used to
34
compute a ratio, F, 2 /F 0 ,. This ratio is then used to obtain thenew value of gross thrust, Fw which is found by the ratio:
FG,.w = FC•--FGFGJ
The ratio procedure is used to minimize any inaccuraciesthat may be caused by assuming burner efficiency (MB) is constantfor all engine operating conditions.
The net thrust and fuel flow after correction for inletrecovery are:
-WV RFFN = FG. ,,-g RFMIL
WFR WF R
and the installed propulsion system thrust and SFC are:
FNA = FN, - DINLET - Dwz+ Dz IMF
SYCA - F,
NOZZLE PROCEDURE
The purpose of the nozzle/afterbody drag and CFG input dataand calculation procedures is to calculate nozzle internal lossesand nozzle/afterbody drag.
NOZZLE/AFTERBODY DRAG
The nozzle/afterbody drag is computed using tables whichrepresent the afterbody drag characteristics (Figure 2.3-6) as afunction of P.,/Po, A9/A10, Mo, external input geometry and enginedata. Parameters obtained from the engine calculations includenozzle throat area, nozzle pressure ratio, freestream conditions,and ideal gross thrust. An essential geometry input is thenozzle exit area, A,, which is required for boattail dragcomputation. This parameter is obtained in one of two ways:
35
T/"MACH, Ps/ P0 AS / A 10
TOI O .AS/ASOR P.S.
CALCULATE AFTBODY DRAG
CO 9/P 0-0.5
AA IA.0A/A 10
MACH
ASM/Al 10-WI-
MACH 4
MAX A/B
6 OR CF IGNRMEDIATE
(b) L r I/ Po (C) PTI / PO
AFTBODY DRAG AND THRUST COEFFICIEN
Figure 2.3-6. Nozzle Performance Calculation
36
1. From the engine input data when the existingaxisymmetric nozzle data are used,
2. From a calculation of fully expanded A, as a function ofnozzle total pressure ratio.
The nozzle/aftbody drag coefficient is shown in Figure 2.3-6(a). The drag coefficient is obtained as a function of theratio of nozzle exit area to maximum cross-sectional area, A,/A1 0 ,and free-stream Mach number, and nozzle exit static pressureratio, P,9/Po. An illustration showing the nozzle aftbody dragprocedure is presented in Figure 2.3-7.
NOZZLE GROSS THRUST COEFFICIENT
The nozzle gross thrust coefficient (CFG) tables are used toprovide a means for correcting uninstalled engine data for theeffects of nozzle internal performance that is different from thenozzle internal performance used in generating the uninstalledengine data. The use of a thrust coefficient table is optional.If no table is used, however, the program will calculate anadjustment to the CFG of the uninstalled data and use this newCFG to find the new installed thrust. The adjustment is onlymade if the nozzle conditions result in over or under expansionlosses.
Two different types of data input formats are provided forthe CFG tables. They are shown in Figures 2.3-6(b) and (c). Thefirst table shows nozzle gross thrust coefficient as a functionof nozzle static pressure ratio and area ratio. Ag/A. iscalculated from tabulated input values provided along the secondtable; however, the nozzle gross thrust coefficient is input as afunction of nozzle total pressure ratio and maximum afterburningand intermediate (dry) power settings. This input data format isbased on the use of a variable area nozzle which is scheduled toprovide an optimum variation of area ratio as a function ofnozzle pressure ratio. The engine power setting and nozzlepressure ratio are obtained from the engine input data in theengine performance calculations.
NOZZLE REFERENCE CONDITIONS
The calculated installed propulsion system performance datainclude the throttle-dependent inlet and nozzle/aftbody losses.To determine the throttle-dependent portion of the nozzle/aftbodydrag to be included as a loss to the propulsion systemperformance, a reference condition has been established for thenozzle/aftbody drag as follows:
37
INTEHNAL INPUTS
ENGINE EXTERNAL INPUTSPROGRAM
DECK AFTBODYGEOMETRYS~AIO
"A1 0 9 REF
ENGINE FLOWPARAMETERS
A 9 T8/P ,A8
8
NOZZLEAFTBODY
DRAG
I ~TABLESI
NOZZLE-AFTBODYSUBPROGRAM
NOZZLE
CFGTABLES
I TOTAL
NOZZLE- RETURNAFTBODY
DRAG AND TO MAINCF PROGRAM
G
Figure 2.3-7. Nozzle-Aftbody Procedure
38
The nozzle/aftbody increment to be included in propulsionsystem installed net thrust will be defined as zero when thenozzle is at its maximum (full-open) geometry and opttratingat a nozzle static pressure ratio, Ps,/Po, equal to 1.0(fully expanded). The nozzle/aftbody drag at this conditionwill be included in the aerodynamic drag. Incrementalchanges in nozzle aftbody drag due to changes innozzle/aftbody geometry and/or nozzle static pressure ratiowill be included as propulsion system drag. This referencecondition is illustrated in Figure 2.3-8 for a typical setof nozzle/aftbody drag data.
THERMODYNAMIC PROPERTIES
Thermodynamic properties required for throat calculationsare obtained using the functions shown in Figure 2.3-9. Thefunctions listed here are "curve-fits" of Keenan and Kaye data.The gas tables are primarily used to calculate exhaust nozzlestatic pressures and jet velocities.
ENERGY BALANCE FOR EXHAUST GAS CALCULATIONS
If the temperature at the engine compressor face, airflow,the bleed mass flow (BL), pressure ratio and fuel flow are known,the exhaust gas enthalpy (h) and relative pressure (Pr) can becalculated from the energy balance:
W2 h + WfOh, = WlahT,. + W+hT + WBLhT
(for either mixed or non mixed flow engines)
For mixed flow fans or a turbojet:
W = W2 - WBr +÷ r
(f/a)a = Wf/(W 2 - Wnd)
hT (Wh T + WfQhB) /We (WjfhBL is considered negligible)
Pr•.. =lF(hT, 0/a)
39
i00
-U---- Go 0 ~ soSi OM
Z* E -4-~ -30d C d E
In~.Y.L o~,d
- ~~o6ea~pubot ' Do
d-- 01 .
140
THERMODYNAMICSUBROUTINE CALCULATIONS
Enthalpy as a function oftemperature (degrees R) and
H = HOFT (T, FOA) fuel-air ratio
Temperature as a function ofT = TOFH (H, FOA) enthalpy and fuel-air ratio
Relative pressure, (Pr) as afunction of enthalpy and
PR = PROFH (H, FOA) fuel/air ratio
Enthalpy as a function ofrelative pressure and fuel-air
H = HOFPR (PR, FOA) ratio
Sonic velocity as a functionof total enthalpy and fuel-air
C = COFH (H, FOA) ratio
Sonic velocity as a functionof static enthalpy and fuel-
C = COFHS (H, FOA) air ratio
Figure 2.3-9. Thermodynamic Subroutines
NOZZLE GROSS THRUST CALCULATION
The calculation procedure in this section applies to bothmixed and non mixed flow nozzles.
CONVERGENT NOZZLE
The velocity at the throat for a convergent nozzle is afunction of the total enthalpy (assuming the throat is choked).
Ce = f (hT, f/a).
and the static pressure is a function of the static enthalpy
hR+ (CR) 217 To = hs +2gJ
To = f(h, f/a)R
Pr, = f(h,f/a)s
41
PT, IPe = (PdrT B/lPý
PTo is obtained from the tabulated engine input data as aF(P.S., alt., M) or it is calculated by the proceduredescribed in the "Nozzle Pressure Ratio Calculation"section.
Pe = PTo/(Pr,/Pe)
The area of the throat isWRT
PCs
and the thrust is
F9, ÷. As* (PS - P,,)Y-g
CONVERGENT-DIVERGENT NOZZLE (Fully Expanded)
If the exhaust flow is fully expanded, the static pressureof the nozzle exit is equal to ambient, and the exit velocity isa function of the total to static enthalpy.
Pr, = P1, (P/,.blP)
h9 = f(Pr, f/a) 9T9 = f(h, f/a)9
Since h. = h9
V9 = [2gJ(hTo, - h 9 ) ]1/ 2
The exit area is
A, = W9R 8 T9 IPmbV
and the gross thrust is
g g
42
CONVERGENT-DIVERGENT NOZZLE (Not Fully Expanded)
If the exit area of a convergent-divergent nozzle is lessthan required for full expansion, the exit static pressure willbe higher than ambient. The throat conditions are known;therefore, a guessed exit velocity gives:
h9 = hT. - V9/2gJ
T9 = f(h, f/a) 9
P,9 = f(h, f/a) 9
P =Pr9Pro
P V9 = (PA V) 9 = (PAV) 9Wg= R(T 9) RT
(C, - stream thrust coefficient)
An iteration on V9 to make W9 = W. will result in the exitconditions for a given area.
The gross thrust is:
Fg = ( WV ÷PA)9 C., - P.A 9g
NOZZLE PRESSURE RATIO CALCULATION
The exhaust nozzle pressure ratio can be calculated ifthrust, fuel flow, and airflow are known. The gross thrust iscalculated as follows:
Pg = (Fnet IF r..) CF,
Fram -g
and the nozzle exit conditions are calculated by assuming thatflow is fully expanded:
43
W = W2 - WDx + WfhlT h, ht= h. + (QfWf/W 8 )TT, = f (hT,f/a)S
V9 =F,7(g) /Wsh9 = hT - V9 /2gJ
P-, = f(h, f/a).(Pr,) 8 = f(hT,f/a)8
since P9 =P--,
The pressure ratio calculation will be in error, an amountrelative to the value of the thrust coefficient (C,.), becausethis is usually unknown if pressure ratios and exhaust areas arenot given.
2.4 Program Structure
2.4.1 Overview
The PWSIM program system consists of:
(i) a driver program that controls the sequence ofcomputations determined by the input optionsselected by the user
(ii) a library of propulsion installation routines
(iii) a set of formatted data files containing propulsioninstallation data for a wide variety of engineinstallation configurations and operating conditions
(iv) a library of mission performance calculation modules
(v) a set of libraries of baseline geometry, drag andweight scaling routines - one library for each ofthe baseline configurations
(vi) an input data set defining the user's selection ofthe various program options and the parametersdescribing the deviation of design from thebaseline. This data set consists of a FORTRANNAMELIST containing both numerical and characterdata
44
(vii) a formatted data set containing definitions of up to20 missions
(viii) a further set of input that is supplied by the usereither interactively (at the computer terminal) oras data entries in the input stream of a batch job.
The complicated nature of engine installation calculationand the large amount of data generated during the installationcalculations necessitates that the program be arranged in anoverlay format to keep the core memory requirements withinacceptable limits. Figure 2.4-1 illustrates the hierarchy of theoverlay structure (and also indicates the names of the routinesthat are accessed within each overlay).
2.4.2 Program Flow
The sequence of activity in the main overlay is shown as afunctional flow chart in Figure 2.4-2.
The sequence of calculations, shown in Figure 2.4-2 is asfollows:
1. Fetch the General Input and Mission Definition Files tothe local operating system.
2. (i) Read and check the MISSION input file(ii) Read the GENERAL input file (NAMELIST file)
3. (a) If an error is detected or an "end of job" inputflag (ENDJOB = 'YES') is read, stop execution;
(b) If all is well proceed to Step 4.
4. Execute the baseline geometry calculations to evaluate:
engine scale
nozzle area
aftbody drag reference area
These data are required for subsequent engine installationcalculations.
5. (a) If installed engine data are already availableproceed to Step 9 (this is denoted by the flag ENGRED set to'YES')
45
cc
0~CI0
LL. wL
owca cc
CJ CC
IL~
100
cc Aý
0
910 0 '-i--
Cq cu0z6 0.
(L46
ACCESS: AIRPLANE DATA FILEMISSION DATA FILEm
FDO GEOMETRY CALCULATIONFOR BASELINE AIRPLANE
A INIAALLED AENIINE DATAREAD?
INSTALLATION
OVERLAY (PWSN., 2,0,0)(• GEOMETRY DRAG WEIGHT
ACCESS: INSTALLED ENGINE DATA CALCULATIONSI
FILE, MARK 11 FILEI(•t OVERLAY (PWSI. 1.0.0)
ENGINE INSTALLATIONCALCULATIONS
TEND
@[ OVERLAY (PWS*M.2.0,0)
GEOMETRY DRAG WEIGHTPERFORMANCE CALCULATIONS
Figure 2.4-2. Functional Flowchart of PWSIM
47
(b) If no engine data have been accessed proceed toStep 6. This is detected by the flag ENGRED set to a value of'NO'.
6. (a) If the engine installation procedure is to be used(if INSTREQ = 'YES') go to Step 8.
(b) If the engine data is to be read from a file ofalready installed engine performance - hereafter referred to as a"Mark 11" file - (if ENGRED = 'NO') go to Step 7.
7. Access the specified Mark 11 installed engine data fileand read its contents into memory. Set the flag that indicatesthat engine performance data has been read (ENGRED = 'YES').
8. Perform the geometry calculations needed to calculatethe aftbody drag reference area, A10, as this quantity is neededfor the installation calculations. On the first pass through thegeometry calculations, the airplane designer's guess at the inletcapture area may be used for the geometry calculations. Thecorrect value is obtained from the installation calculationswhich cannot be performed until A10 is known. Thus, a two-stepiteration is necessary. The first iteration calculates the A10value using an assumed capture area and then recalculates thecapture area using the new A10.
9. Perform the engine installation calculations using thecurrent value of A10, calculate the correct value of referencecapture area, RACAPT. If this is the second time through thisblock:
write results to TAPE20
set flap to show installation is complete (INSTREQ = 'NO'
set flap to show engine data has been read (ENGRED = 'YES').
10. Check to see if the installation calculation hasalready been accessed.
a. Return to geometry calculation with correct valueof RACAPT
b. Go on to Step 11.
11. Perform the airframe technology calculations (geometry,drag and weights) and the airplane mission calculations. Writethe results to TAPE 6.
48
12. Check the status of the 'end of job flag' ENDJOB
a. If ENDJOB = 'YES' stop execution
b. If ENDJOB = 'NO' go to Stop 2-(ii).
It is to be noted that the engine performance data areaccessed (by reading the MARK 11 file or by installation) onlyonce per job. If subsequently the airplane size (and thus enginesize) is changed (for example, during a sizing iteration), thenthe installed engine performance data are scaled rather thanperforming another installation. To be strict, the uninstalleddata should be scaled prior to installation; since airplanesizing can involve several iterations of engine size, thetechnique of scaling the installed data is used to keepcomputation time as low as possible.
An alternative way of looking at the program overallstructure is shown in Figure 2.4-3. This shows the main programmodule and the three subservient libraries. The library showninside the box of dashed lines is that which contains thegeometry, drag, and weight modules for the configuration beingstudied.
3.0 Data Base Descriptions
Seven weapon system preliminary designs are provided toserve as point-of-departure baseline configurations for the PWSIMprogram.
The conceptual data bases produced for each configurationconsist of several items that, taken together, will give athorough definition of the design-point aircraft and provide asound basis for parametric studies. The items contained in eachdata base are:
(a) an outboard profile, three-view engineering drawing
(b) engineering description
(c) geometric summary
(d) weight statement
(e) drag polars
(f) engine performance data
(g) airplane performance in the design mission
(h) limitations on the applicability of the data base.
49
r -- - - - -
0 0
wr i -a -
w a.w CO L
> LU 0
I-
CLL
500
The following sections contain detailed descriptions andcomprehensive data summaries of the selected baseline aircraftconfigurations. Subsequent sections contain representative dataallowing the evaluation of the drag and weight penalties ofexternally carried stores.
3.1 Tactical Fighter - Model 985-420
3.1.1 Concept Description
This aircraft is shown in Figure 3-1. The overall airplanelength is 62 ft 2 in and the wing span 50 ft 9 in. The wing hasa leading edge sweep of 37.5O, a reference wing area of 571 ft 2,an aspect ratio of 4.5, and a wing thickness ratio of 5% at theside of body and 4% at the tip. A smooth variable camber leadingedge is used with a hinged, single slotted trailing edge flapduring landing approach. Wing camber is varied automaticallythroughout the flight envelope for improved lift/drag ratio.Hardpoints are provided for carrying external fuel tanks (forextended range and ferry missions) and alternate weaponconfigurations.
The airplane is designed for a one-man crew. Locatedforward, aft, and below the crew compartment are theavionics/electronics equipment compartments. Included in the1859 lb of avionics equipment are target acquisition,communication, navigation and identification, informationmanagements, and defense functions. ECS equipment, oxygen, andelectrical/hydraulic subsystem equipment are located in thefuselage aft of the pilot. The body fuel is carried in integraltanks with a capacity of 12,000 lb of JP-4 fuel.
Two vertical fins are integral with the aft fuselage sidewalls and have a total area of 110 ft 2 . Each uses a conventionalrudder (32% of the fin chord). All-moving, slab canards with anexposed total area of 78 ft2 are used for longitudinal and rollcontrol throughout the entire speed regime. Wing flaperons willaugment roll control throughout the flight envelope.
3.1.2 Aerodynamics
Estimated aerodynamic characteristics are presented in thissection for the 985-420. Figure 3-2 illustrates the completedrag polar at three key conditions in the flight envelope.
3.1.3 Weights
The weight statement for the 985-420 is shown in Figure 3-3.Weight estimating ground rules and assumptions are:
51
* I I
ol
s t
.41
'-I
'-4c
"44
ff0
'u, IV f
U9 x
52
1I 1 MODEL 985-420
080.9/30K __ -.
10.6 1. 2/45K
c 1 0.4 I - -- - 2~.06K -
___ ~CRUISE -
10.2 i t,
- n5 0 .1 0 05, 010
Figure 3-2. Tactical Fighter, Drag Polars
53
985-42 NOSE STATION 0.GROUP WEIGHT STATEMENT WEIGHT-LBS WING MAC 151.PDWTS 01-OCT-82 VERSION LEMAC 407.3-MAY-83 BODY LENGTH 746.
Body St Percent MACWing 3728. 483.Canard 566. 295.Vertical Tail 665. 678.Body 5180. 411.Alighting Gear 1541. 410.Nacelle or Eng Section 188. 603.Air Induction System 963. 341.
Total Structure 12836. 438.Engine + Accessories 4960. 603.Starting + Control 160. 387.Fuel System 709. 420.
"o Fuel tanks are inerted with nitrogen gas from N2
generators.
"o Inflight refueling provisions have been incorporated.
"o Arresting tail hooks have not been incorporated.
"o Engine inlets, pitot tubes and canopy have de-icingsystems.
"o Air conditioning systems are closed loop bootstrap plusliquid cycle cooling variety.
"o Provisions for weapon hardpoints on wing and fuselagehave been incorporated. Each hardpoint assumes multiweapons control in terms of attachment and launching.
"o The APU is an IPU "Integrated Power Unit." Its functionis to operate as a starter and an emergency power system.
"o The landing gear CBR is 9. Note: CBR (CaliforniaBearing Ratio) is a measure of the bearing strength ofthe airfield from which the aircraft must operate.
"o Hydraulic system operates at a pressure of 4000 psi.
"o Flight control system utilizes fly-by-wire technology.
"o Flight design weight equals gross weight less 20 percentof the on-board fuel weight.
"o Landing weight equals the gross weight less 40 percent ofthe on-board fuel weight.
"o No weight penalty has been assessed for incorporation ofMission Adaptive Wing (MAW).
"o TAD is 1987, IOC is 1993.
3.1.4 Propulsion System
Uninstalled engine data were computed using the PWA enginecycle program, PWA CCD 1178-06.01. This engine has a bypassratio of 1.0, an operating pressure ratio of 25, and a max burnertemperature of 3000'F. Installation effects were estimated usingthe Engine Installation Analysis Program. The inlets are underwing, centerline mounted, two-dimensional external compressiondownward spilling fixed horizontal ramp, which allow the airplaneto achieve the M = 2.0 dash speed at altitude, and provide inlet
55
flow protection during high angle-of-attack maneuver conditions.The inlet ducts are designed with structural radar-absorbentmaterials (RAM). The capture area of each inlet is 5 ft 2 .
Engine mounted 2-D, C-D nozzles are arranged side-by-sideand incorporate variable throat area capability for augmentedengine operation. Adequate cooling flow is provided to nozzlesurfaces to minimize IR signature and to allow application of RAMfor reduced RCS. Installed thrust and SFC curves for subsonicand supersonic cruise conditions are shown in Figure 3-4 through3-5.
3.1.5 Performance
This aircraft was designed to fly to a radius of 1000 nmiand patrol on station for 1 hour before returning to the startingpoint. A summary of the basic sizing mission is shown in Figure3-6. A summary of the design mission segment by segment is givenin Figure 3-7.
3.2 Supersonic Interceptor - Model 985-430
3.2.1 Concept Description
This vehicle, illustrated in Figure 3-8 has an overallairplane length of 93 ft 4 in and a wing span of 38 ft 5 in. Thewing has a leading edge sweep of 750 on the main inner wingsection and 550 on the outboard section, a reference wing area of1002 ft2 , an aspect ratio of 1.47, and a wing thickness ratio of4.4% at the side of body and 1.9% at the tip. A smooth variablecamber leading edge is used with a hinged, single slottedtrailing edge flap during landing approach. Wing camber isvaried automatically throughout the flight envelope for improvedlift/drag ratio. At low speeds, the leading edge vortex flap isdeployed, as is the high lift canard. The wing provides volumefor approximately 7770 lb of fuel in integral wing tanks.
The airplane is designed for a one-man crew. Locatedforward, aft, and below the crew compartment are theavionics/electronics equipment compartments. Included in the2699 lb of avionics equipment are target acquisition,communication, navigation and identification, informationmanagement, and defense functions. ECS equipment, oxygen, andelectrical/hydraulic subsystem equipment are located in thefuselage aft of the pilot. The body fuel is carried in integraltanks with a capacity of 18,130 lb of JP-4 fuel.
The vertical fin has an area of 130 ft 2 . A conventional 30%rudder is incorporated. Additional directional stability isprovided by a ventral fin.
56
CDCD
' II 1i-
CD
00 1
I I Co
-o O
cn 00
C -M " \ .- , 4
C4i :: w "
2 -H-E-W~
, C
zoli
57\ Il II J I •4
I_ II ' i I U
I I I I I J I
I
11111111 • L '' i I I j-
57
I C)
I 00
0In1
000
CCýl -0 '.0C4 C14 .
I4J 58
cc
U. E-
0 = L0
w C U
a~ U. _ = au
- us~' 4 LU(8 0LS2u- -X CC 31- < 8 co W
us 4 (c W .
U. ~ ~ UJ W i j
0A14 0C 000,
5 9
0 0 tv) 0 O0
H v- NH
U) C) 0q C)
0o ( L) ) 0f 0
0-.r40)
0 r0
C>~ ~ ~ M D0 D
H N A..
0)
41
Co0 0 C)0 0
H H N 0 C) h0 Y Cop 0.
q qT (1) H , Uu
4-)
E-4
N Ný0 0; 40 a 0 N)
0 5.4jJ() 0&
hir
Hhi
60 0
.00,IlIk
-4.-6
af
__
.61
- 4 I'
I I•• •
0 2!
The stowable canard has an area of 52 ft 2 and provides ahigh lift capability at low speed through the use of aslat/double slotted flap airfoil. Roll control is provided bywing spoiler/slot deflectors and trailing edge flaperons.
3.2.2 Aerodynamics
Estimated aerodynamic characteristics are presented in thissection. Drag polars for the critical mission Mach numbers areshown in Figure 3-9.
3.2.3 Weights
The weight statement and weight related design data aretabulated in Figure 3-10. Weight estimating ground rules are thesame as those applicable to the tactical fighter (985-420) andare listed in Section 3.1.3.
3.2.4 Propulsion System
The design mission for the Model 985-430 requires an enginethat can operate with low specific fuel consumption during cruise-at Mach 3.0 and an altitude of 70,000 feet. To meet this goal,engines of bypass ratio of 0.2, overall pressure ratio of 10 anda maximum burner temperature of 3000'F were selected.
Two General Electric Mach 3.0 advanced technologyafterburning (GE16/J6-Bl) and dry (GE16/J5-H3R) turbojets providethe necessary propulsion. The inlets are under wing,axisymmetric mixed compression, which allow the airplane toachieve the M 3.0 combat speed at altitude, and provide favorableinterference with the wing. The inlet ducts are designed withstructural radar-absorbent materials (RAM). The capture area ofeach inlet is 10.2 ft 2 . Engine mounted axisymmetric nozzlesincorporate variable throat area capability for augmented engineoperation.
Installed SFC data are presented in Figures 3-11 and 3-12.
3.2.5 Performance
The Supersonic Interceptor has been designed to fly a 1000-nautical-mile radius intercept mission out and back at Mach 3.0(Figure 3-13).
A summary of the design mission history is shown in Figure3-14.
3.3 Supersonic Intercontinental Cruise Missile
62
-~- 0
-00
(U.'
-"0
'o 6T
C-4-
'too
----- T)
o 03
985-430 INTERCEPTOR NOSE STATION 0. INGROUP WEIGHT STATEMENT WEIGHT-LBS WING MAC 376. INPDWTS 01-OCT-82 VERSION LEMAC 503. IN9-MAY-83 BODY LENGTH 1120. IN
Body Sta Percent MACWing 5420. 678.Canard 220. 197.Vertical Tail 801. 980.Body 4559. 504.Alighting Gear 2422. 597.Nacelle or Eng Section 703. 865.Air Inducting System 483. 768.
Total Structure 14607. 631.Engine + Accessories 6342. 865.Starting + Control 150. 768.Fuel System 949. 670.
An outboard profile of this vehicle is shown in Figure 3-15.The overall vehicle length is 30 feet and the wing span is 15.5feet. The wing has a leading edge sweep back of 70 degrees and atrailing edge sweep forward of 40 degrees. The wing has aconstant thickness/chord ratio of 0.03 and a taper ratio of zero(except that rounding the wing tips preserves a finite materialthickness at the tip).
The design is sized to achieve a range of approximately 3500nautical miles using present state-of-the-art turbojet propulsionand JP-10 type fuel. The payload consists of a single ballisticvehicle having a suitable yield. The ballistic vehicle is shapedand treated with radar absorbing material for penetration of theterminal defenses in the target area. The avionics systemincorporates an inertial system having the capability to receiveupdates from a Star Tracker or GPS thereby achieving the requiredterminal accuracy.
The reference midcourse cruise configuration is designed forair launch from a carrier aircraft at a Mach number of 0.6 at30,000 feet altitude. Solid rocket boosters take the missilefrom air carrier loiter conditions of M = 0.6 and 30,000-footaltitude to cruise Mach number and altitude of 3.5 and 85,000 ft,respectively. Two boost motors, one on each side of the lowersurface of the fuselage/wing intersection, are used to minimizeoverall carriage length.
Insulated structure is employed with the insulationprotected by a thin outer layer of titanium having a highemissivity coating. The load carrying structure is Epoxy-Graphite for those parts of the structure where temperatures donot exceed 400F or Polyimide Graphite (up to 6000). The min kinsulation passively cools the fuel so that a 3500 condition(with 35 PSIA vapor pressure) is not exceeded for the flight.Launch is assumed at high altitude, -65*F condition. The warheadis also passively cooled. An allowance for active cooling of theelectronics is included in the fixed equipment.
3.3.2 Aerodynamics
The drag of this configuration has been estimated using thpresults of wind tunnel tests carried out in the NASA Ames A- x 2-foot transonic and 10 x 14-inch supersonic wind tunnels on amodel approximating this configuration. The model differed fromthe current design in that:
o the tested model had a semicircular, underwing fuselage
69
S~i°.
II..•-•-•I.. .
700
nmnnnmnmuunuulnnn numuumum Ki mini
"o the tested model had a rearward sweep wing trailing edge
"o the tested model had no engine installation
"o the tested model had one centrally located vertical finwhereas the current design has two fins mounted at the65% spanwise station on the wings
"o the test Reynolds Number at M - 3.5 was 5.9 millioncompared with 22.0 million for the full-scale vehicle.
Appropriate corrections to the test data result in the drag
polars shown in Figure 3-16.
3.3.3 Weights
The weight statement for the baseline Supersonic CruiseMissile is shown in Figure 3-17.
The weight calculations take due account of the designpeculiarities of this vehicle (with reference to conventionalairplane design methods). Confidence in the approach taken hasbeen enhanced by using the Boeing weights methodology tocalculate the weight of the BAC ALCM"B" - a configuration forwhich detailed weights data are available.
Design considerations influencing the weight calculationinclude:
o Airframe Construction
Wing - The wing is constructed with a center core coveredwith a 0.2-inch-thick skin of polyimide/graphite material.Forward and aft of the center core are sections of chord 20inches that have a similar structure but with a thin (0.05-inch)skin of titanium bonded to it. Forward and aft of this regionare 14-inch chord sections constructed of a honeycomb core with a0.08-inch titanium skin. The leading and trailing edges and wingtip are made of solid titanium.
Fuselage - The fuselage is of skin-frame constructioncomposed of polyimide/graphite material with an outer skin of0.02-inch gauge titanium. Between the polyimide/graphite andtitanium skins in a 0.125-inch layer of insulation; this keepsthe fuel temperature below 350'F.
o Radar absorbing material is applied at appropriate partsof the airframe
FLIGHT CONTROL 233.HYDRAULIC, PNEUMATIC + ELECTRIC 173.AIR COND + ANTI-ICING 97.LOAD + HANDLING 10.
TOTAL FIXED EQUIPMENT 513.
WEIGHT EMPTY 4801.
OIL + UNUSABLE FUEL 42.
NON-EXP USEFUL LOAD 42.
OPERATING WEIGHT 4844.
PAYLOAD 400.FUEL 3356.
GROSS WEIGHT 8600.
Figure 3-17. Cruise Missile Weight Statement
7 1
"o No crew or crew accommodation equipment is included
"o No landing gear is present
"o High density fuel (JP-10 synthetic, hydrocarbon) isemployed (no ullage allowance is included in the fueltanks; a bellows arrangement allows for fuel expansion)
"o Load factor was selected to be 12 to allow for safe
carriage by appropriate aircraft.
3.3.4 Propulsion
The supersonic intercontinental cruise missile was designedto use a high temperature, nonaugmented, turbojet engine. Thecruise missile has been designed with a Mach 3.5 two-dimensional,mixed compression inlet system. The inlet features an initialcompression ramp of 7'. A variable ramp system was used toprovide efficient external compression at the design conditionand low spillage drag at off-design conditions. Porous boundarylayer bleed surfaces were located on all four sides of theinternal duct. The bleed was passed into three compartmentedbleed plenums and exhausted overboard. A bypass system was alsoincluded for engine inlet matching and to enhance inlet restartcapability.
A fixed geometry, expansion/deflection nozzle was selectedfor the cruise missile to reduce the overall engine installationlength. The nozzle was designed to use a Prandtl-Meyer expansionthat is formed from the nozzle throat to the exit plane about abase plug. The resulting supersonic contour was short, so thatfrictional losses were lower than for conventional nozzles. Thisgain was offset, however, by the drag due to low pressure on thebase plug.
It is important to note that the cruise missile has beendesigned for a Mach 3.5, 85,000-ft supersonic cruise conditionand that the fixed area expansion/deflection nozzle has beendrawn for a minimum power setting. Cruise thrust and SFC areshown in Figures 3-18 and 3-19.
3.3.5 Performance
The portion of the cruise missile mission that employs gasturbine propulsion is shown in Figure 3-20. The gas turbine isstarted at a Mach number of 3.5 at 85,000 feet altitude.
The mission consists of a cruise-climb to about 98,000 feetat constant Mach number. The mission ends at the point where the
74
3000
85000 FT
2000
NET THRUST
94000 FT
1000
00 1.0 2.0 3.0 4.0
MACH NUMBER
Figure 3-18. Cruise Missile Thrust Available
7r
4.0
Altitude:85000-94 00 Ft
3.0
SPECIFIC FUEL M-3.0CONSUMPTION
(LBAHRII/LB
1.0
0 20 40 60 80 100 120
CORRECTED THRUST (FN) - 1000 LB
Figure 3-19. Cruise Missile Specific Fuel Consumption
76
10OOAO
95.0oo
ALTITUDE"-. FT
90,000
J - Cruise - limb
- M=3. 5
85,000
0 002000 3000 4000
RANGE -NMi
Figure 3-20. Cruise Missile Design Mission
77
fuel is all expended; at this point, the ballistic payload isreleased over the target area. The airframe is not recovered.
The cruise is carried out at optimum lift-drag ratio and ata constant power setting (maximum dry thrust). The resultingrange factor varies from about 6900 to 7100 nautical miles overthe extent of the cruise.
3.4 Long Range Transport - Model 1046-103
3.4.1 Concept Description
This aircraft is shown in the three-view engineering drawingin Figure 3-21.
The aircraft has an overall length of 220.8 feet andwingspan of 206.84 feet. The wing has a leading edge sweepbackangle of 370, a reference area of 4754 square feet, an aspectratio of 9, and wing thickness-to-chord ratio that varies from0.12 at the root to 0.08 at the tip. High lift for takeoff andlanding is provided by full-span leading edge slot and double-slotted trailing edge flaps.
The body is designed to carry a payload of 200,000 poundsconsisting of heavy and/or outsized cargo. The cargo compartmentis 142.8 feet long, 17.5 feet wide, and has a maximum height of13.5 feet. The body has cargo doors and a loading ramp under theupswept rear fuselage and a hinged nose thus providing a drive-through capability. The high-flotation landing gear (withkneeling capability for easy loading) is housed in pods locatedon the lower part of the fuselage.
The fuselage volume is totally dedicated to cargo so no fuelis carried there. All the fuel is stored in the wing.
The aircraft has conventional, horizontal and vertical tailsof area 1060 and 786 square feet, respectively. Elevators andrudder of 30% chord provide flight control surfaces.
Propulsion is provided by four nacelle-housed P&W parametricturbofan engines of bypass ratio 5.74 sized to produce 30,050pounds of static thrust.
The design constraints imposed on this configuration
include:
"o Payload 200,000 pounds
"o Range 4,600 nautical miles
78
C144
CD
... I. CC)
* D0 '4
I *, (\J('
000
(.4
In Ila0 - C" 4. z
U. co
79
o Critical Field Length 8,000 feet
o Load Factor 2.5
"o R/C with OEI 100 feet/minute
"o No fuel stored in fuselage
3.4.2 Aerodynamics
The aerodynamic characteristics of the Model 1046-103 in theform of drag polars for important Mach numbers are shown inFigure 3-22.
3.4.3 Weights
Figure 3-23 shows the weight statement of the Model 1046-105. Weight estimation is consistent with a TAD of 1990.
3.4.4 Propulsion System
The nonaugmented turbofan engine selected for the Model1046-103 transport concept was chosen engines investigated in theAdvanced Technology Engine Studies (ATES) program. The enginecycle included a bypass ratio of 5.74, an overall pressure ratioof 35.0, and a combustor exit temperature of 2600'F.
A turboprop engine, the Pratt & Whitney STS679, has alsobeen supplied for use with the Model 1046-105. This three-spooladvanced technology engine features a two-axial stage, onecentrifugal stage, high compressor driven by a single-axial stageturbine, a four-axial stage low compressor driven by a single-stage turbine, and a gearbox driven by a three-stage freeturbine. The overall pressure ratio of the engine is 27.5, thecombustor exit temperature was 2379°F, and the speed of the powerturbine is 10,960 RPM.
The propeller selected for use with the STS679 was chosenfrom a previous Boeing in-house study of near-term and advancedpropellers supplied by Hamilton Standard. Propeller tip speedand loading were also selected based on this study. The systemchosen was a counter rotating prop fan. This small diameter,highly loaded, multibladed, variable pitch, unducted fan has beendesigned for use on aircraft with cruise speeds up to Mach 0.85.
Thrust and SFC of the installed turbofan engine are shown inFigures 3-24 and 3-25, respectively.
80
1.0
M.0.3-ALT,, SLLOITER
0.8
0.6-
LIFT CRUISE,COEFFICIENT C 40,000 FTCL M-,0.75
CL
0.4
0.2
0 0.02 0.04 0.06 0.08 0.10
DRAG COEFFICIENT, CD
Figure 3-22. Long Range Transport Drag Polars
81
Group Weight Statement Weight Design DataIADSWT2 07/01/77 Version Lbs
SEXP 3840Engine + Accessories 22507 Aspect Ratio 9.00Thrust Reversers 3117 Taper Ratio .36Exhaust + Deflectors 0 TOC Root .12Fuel System 2317 TOC Tip .08Engine Control 100 Sweep E.A. 33Starting System 400 H Tail
SGross 1061Total Propulsion 28440 SEXP 921
Aspect Ratio 4.50Flight Control 7423 Taper Ratio .36Auxiliary Power Plant 931 TOC Root .09Instruments 860 TOC Tip .09Hydraulic + Pneumatic 2139 Sweep E.A. 29Electrical 3528 Tail Arm 112Avionics 3451 V TailArmament 0 SGross 736Furnishings + Equip 4483 Aspect Ratio 1.65Air Cond + Anti-Icing 3103 Taper Ratio .36Photographic 0 TOC Root .10Load + Handling 0 TOC Tip .10
Sweep E.A. 25Total Fixed Equipment 25917 Tail Arm 102
Tank Volume 28970Payload 200000 SystemsPassengers + Baggage 0 KVA Reqd 202Fuel 191770 Volume Pres 50690
Gross Weight 623360
Figure 3-23. Long Range Transport Weight Statement
82
40
30
NET THRUST1,000 LB
10
1040,000
50,000
0 0.2 0.4 0.6 0.8 1.0MACH NUMBER
Figure 3-24. Long Range Transport, Cruise Thrust
83
1.4
1.2
ALTITUDE 36,089 FT
1.0
1.05 (SFC)[- (LB/HR)/LB
0.8
0.6 .8
MACH NO.
0.1
10 20 30 40 50
CORRECTED THRUST, FN/ 6 - 1.000 LB
Figure 3.25. Long Range Transport, Cruise SFC
84
3.4.5 Performance
The Model 1046-103 was designed to carry a payload of200,000 pounds of cargo over a range of 4600 nautical miles. Tominimize aircraft size and cost the mission is flown at optimumaltitude and cruise Mach number. The mission is illustrated inFigure 3-26.
A summary of the designed mission is shown in Figure 3-27.
3.5 Lightweight Fighter - Model 985-213 (Modified)
3.5.1 Concept Description
The vehicle consists of a blended wing-body configurationwith twin vertical tails mounted on the wings at about the 3/4span location (Figure 3-28).
The overall length of the aircraft is 44 feet and thewingspan is 19.7 feet. The wing has a NASA SCAT 15 plan formwith 740 leading edge sweep, with an aspect ratio of 1.46, taperof 0.19 and a reference area of 266 square feet. The wingthickness varies from 4% at the root to 3% at the tip. Wingcamber is variable throughout the flight envelope.
The aircraft carries a one-person crew in a low-profilecockpit at the design takeoff gross weight of 12,500 pounds, thedesign wing loading is 47 pounds per square foot, and thethrust/weight ratio if 1.32.
Wing structure is skin and multispar construction ofgraphite composite material. The structure is designed for aload factor of 7.33 g's at the flight weight of 12,500 pounds, adynamic pressure placard of 2133 pounds per square foot (Mach 1.2at sea level) and a Mach 2.2 dash capability at altitude.
Air-to-air weapon capability consists of two lightweight(CLAW) missiles mounted semisubmerged on the upper aft fuselage.
A 20-mm gun and 250 rounds of ammunition are carried internally.
3.5.2 Aerodynamic
Estimated aerodynamic characteristics of the unmodifiedModel 985-213 are presented in Figure 3-29 through 3-31. Figure3-29 shows the detailed breakdown of drag-at-zero-lift for threeflight conditions.
Trimmed drag polars for typical subsonic and supersonicflight conditions are shown in Figures 3-30 and 3-31,
85
AIRLIFT COMBAT MISSION
4,600 NMI
Q WARMUIP/TAKEOFF - 5 MIN.
( CLIMB
( CRUISE - OPT MACH/ALTITUDE
( LOITER - 30 MIN.Figure 3-26. Long Range Transport Design Mission Profile
respectively. The drag of the new aircraft should be close tothat of the original; differences due to the aft-body referencedrag changes are to be expected because of the differentinstallation.
3.5.3 Weights
The weight statement for the 985-213 is shown in Figure 3-32. Weight estimation ground rules and assumptions are listedbelow:
"o The majority of aircraft structure is advanced composites(graphite-epoxy)
"o Airframe Integrated Nozzle
"o Fly-by-wire surface controls
"o Avionics equipment in compliance with statement-of-workrequirements
"o Semisubmerged CLAW missiles (2)
"o Final aircraft geometry is the result of aerodynamic andweight parametric trade studies and represents the bestcompromise for overall performance
"o Lightweight M-197 20-mm Gatling gun with gas drive
"o Judicious location of gun, ammunition, missiles, and fuelsuch as to minimize CG gravel as these items are expended
"o Fuel pumping for trim control.
3.5.4 Propulsion
Uninstalled engine performance was computed using the Pratt& Whitney Aircraft parametric engine cycle deck, CCD 1178-08.00.The engine is a minimum bypass ratio, dry turbofan having a maxdry uninstalled thrust of 16,500 lb sea level static. The enginecycle characteristics are bypass ratio (BPR) = 0.2, overallpressure ratio (OPR) = 26, turbine inlet temperature (TIT) =30000 F.
The inlet is located under the fuselage, centerline mounted.it is a two-dimensional, external compression inlet utilizing avariable ramp, four-shock system.
This inlet has two movable external ramps, a 7.30 initialramp angle, a boundary layer control bleed system consisting of
92
WEIGHTLBS.
WING 1180HORIZONTAL TAILVERTICAL TAIL 100BODY 1520MAIN GEAR 380NOSE GEAR 110LAUNCH AND RECOVERY GEARENG SECTION OR NACELLE 360
STRUCTURE (3650)ENGINE AND EXHAUST 2200THRUST REVERSERENGINE ACCESSORIES 50ENGINE CONTROLS 80STARTING SYSTEM 100FUEL SYSTEM 340
BASIC MISS FLT DES WT 10,400FULL INTERNAL FUEL 3,630
Figure 3-32. Weight Statement
93
porous bleed on the second and third ramp surfaces, sideplates,and throat bleed slot located aft of the normal shock. Thethroat slot also acts as a bypass to remove excess inlet airflowfor matching engine airflow demand with inlet supply. The inletcapture area is 4.44 ft 2 .
An engine mounted 2-D/C-D nozzle which incorporates a fixedthroat and a variable exit area was utilized for efficient engineoperation.
3.5.5 Performance
The aircraft was configured to provide low drag at thedesign Mach number of 1.8. A design mission was specified (seeFigure 3-33) that involved flight at altitudes limited to 50,000feet (a pressure suit limit). Mission characteristics aresummarized in Figure 3-34.
3.6 Carrier Air Vehicle/Transatmospheric Vehicle
3.6.1 Concept Description
The Model 896-111 is a two-stage-to-orbit system with bothstages being recoverable (Figure 3-35).
The orbiting vehicle is carried in a cavity in the undersideof the fuselage of the first stage. This concept minimizes therequirement for the large amount of ground-support equipmentnormally associated with today's conventional vertical takeoffrocket launch system. The proposed system utilizes a horizontaltakeoff and landing mode.
For mating, the booster and orbiter are each towed to an"alert pad" and the vehicles aligned with their longitudinalcenterlines coincident with each other. The orbiter is thentowed forward into the booster body cavity and mechanicallyjoined to the booster. The orbiter landing gears are retracted,and the booster orbiter combination is towed to the LOX/LH2servicing facility which is adjacent to the TAV pad to allow allcryogenic loading and replenishment to be controlled in one area.After completion of the takeoff, climb, and separation, thebooster would return to the base to be recycled for any necessarymaintenance.
The CAV is illustrated in Figure 3-36. The two-man crew andaircraft subsystems are located in the forward body. The twocylindrical LH2 fuel tanks are paired in the forward fuselagewith the LOX tank pair located directly to the rear. The noselanding gear is located forward and below the LH2 tankage. The
nacelles, with fixed supersonic inlets, are located outboard ofthe body cavity which accommodates the orbiter. The eightwheeled main landing gear is integral with the nacelle andretracts forward into the lower nacelle when stowed. The wingsare mounted high on the fuselage to provide clearance with theunderslung orbiter. Wing tip mounted verticals are used toprovide directional stability. A single SSME rocket engine isused during boost phase and is located on aircraft centerline atthe wing trailing edge.
The booster forward body contains LOX/LH2 rocket propellantsand propellant crossfeed system to the orbiter to ensure that theorbiter vehicle propellant tanks are completely filled at stageseparation. The JP-4 airbreathing fuel is contained in theoutboard wings to reduce the total wing bending moments at theside of body. The booster is designed for a two-man crew.Located forward, aft, and below the crew compartment are theavionics/electronics equipment compartments. ECS equipment,oxygen, and electrical/hydraulic subsystem equipment are locatedin the fuselage aft of the pilot.
The first stage (booster) utilizes present day state-of-the-art construction.
BODY
Body structure is semimonocoque with frame supportedgraphite/polyimide honeycomb sandwich skin panels. Two deepaluminum honeycomb beams form the sidewalls of the orbiterrecess, carrying twin lower-body longerons and providing verticalshear capability. Attached to the wing by the wing-to-bodylongeron, these beams extend aft of the wing and form the inboardstructure of the airbreathing engines mounting structure. Withinthe body cavity, the beams carry the pair of trapezes whichcontrol the relative movement of the booster and orbiter toensure clean separation.
The other engine supports are provided by vertical beamsattached below the wing, the center one acting as a duct splitterover its forward portion, the outboard one forming the nacellewall. Further structure is provided by the horizontal ductsplitter, which continues aft as a firewall separating the upperand lower engine pairs, providing lateral shear stiffness.Engine removal is effected through individual hatches on the topand bottom surfaces of the nacelle. Removal of any or all of thehatches does not affect the structural integrity of the enginesupport structure.
99
The cylindrical LH2 tanks are paired in the forwardfuselage, and are link-supported inside the body monocoque. Foreand aft loads are taken by a thrust structure joining the afttank ring to a body bulkhead which serves to separate the fueland oxidizer bays, and also forms a manufacturing joint. Aft ofthis is the LOX tank pair, also link supported, with a thruststructure to the front spar of the wing. Forward of the LH2tanks are the nose landing gear bulkhead, the equipment and ECSbays, and the crew compartment and capsule.
WING
The high-mounted wing carries the four orbiter attachmentpoints, two each on the front and rear spar center sections. Thefour-spar wing has graphite/polyimide honeycomb sandwich skinswith integral spar caps. Stringers, spar webs, and ribs aregraphite/epoxy co-cured. The wing leading edge is a built-uptitanium structure with provisions for thermal stress relief.Control surfaces are of graphite epoxy honeycomb.
VERTICAL TAILS
The vertical tails are of similar construction to the wing.The possibility of using split rudders is being studied. Thiswill enhance directional stability in slip-flow conditions byforming wedge-type vertical tail surfaces.
LANDING GEAR
The main landing gear comprises two struts, each carrying aneight wheeled truck, retracting forward into the nacelle lowersurface. Vertical loads are reacted to the wing structure by abulkhead spanning between the inboard beams and the outboardnacelle wall.
The nose landing gear is mounted on the bulkhead ahead ofthe LH2 tanks, and retracts rearward to lie below the tanks.Provision is made for emergency extension should the hydraulicsystem fail. Because of the wide spread between takeoff andlanding weights, an Adaptive two-stage oleo design is proposedfor all three elements of the tricycle landing gear.
3.6.2 Aerodynamics
The drag data shown in Figure 3-37 and 3-38 have beenevaluated using several Boeing programs to calculate drag fromvarious sources (skin friction, wave drag, etc.). The data baseestimation uses simplified methods that have been calibrated tomatch the results from the detailed drag analysis.
a. drag coefficient at zero lift as a function of altitudeand Mach number
b. drag coefficient due to lift as a function of liftcoefficient and Mach number
c. 'drag-area' or D/q increments as a function of Machnumber
The latter tables allow the mission analysis program to takeaccount of the drag changes that result when:
a. the rocket engines are fired (drag change due to reducedbase area)
and
b. the TAV is not attached to the CAV (drag change is dueto modified base area and wetted area).
3.6.3 Weights
The weight of the Model 896-111 was estimated using theBoeing Level-i weight estimating program, PDWTS, for theconventional airplane components; rocket engine, cryogenicsystems, etc. were evaluated using detailed analysis of thesystems.
A typical weight statement is shown.in Figure 3-39.
3.6.4 Propulsion System
The first-stage booster is powered by eight advancedaugmented airbreathing engines (F-101 uprated) each producing35,000 lb static sea level thrust and one SSME rocket engine(A*/Ae - 150) having a vacuum thrust rating of 530,200 lb and an
ISPVAC - 463.5 sec using LOX/LH2 propellants. The booster launchsystem utilizes airbreathing propulsion during the takeoff andclimb to 30,000 ft and M = 0.86. At this time, the rocketengines on both stages ignite and operate until reaching 117,500ft altitude and 3000 ft/s velocity where stage separation occurs.
The airbreathing propulsion system performance in themission analysis program is calculated from tables of installedthrust, fuel flow, and corrected airflow of the engines. Theinstalled performance data are calculated by the program using
103
Model 896-111GROUP WEIGHT STATEMENT WEIGHT-LBSWTS 01-SEP-84 VERSION11-FEB-86
manufacturer's uninstalled performance data together with usersupplied inlet, nozzle, and aftbody drag data.
Rocket engine performance is estimated using vacuum thrustand specific impulse corrected for ambient pressure effects.
3.6.5 Performance
The typical mission for the CAV consists of takeoff (with aground roll of about 10,000 feet) and a climb to 30,000 ft and M= 0.862 using augmented, airbreathing engines.
After climbing to 30,000 ft and M = 0.862 under augmentedairbreathing power, all rocket engines are ignited with thetakeoff and climb taking 820.9 seconds. The vehicles proceedthrough a dual burn accelerated climb to the separationconditions under airbreathing and rocket thrust. During thisinitial boost the maximum dynamic pressure experienced is 1050PSF and occurs at an altitude of 40,300 ft at M = 1.97. Thevehicles separate at 117,500 ft, V = 3000 FPS where the dynamicpressure is 65 PSF. The orbiter proceeds to the requiredinjection conditions for the particular mission with itspropellant tanks full at separation. After separation thebooster is lofted to 156,000 ft by its own momentum, descends,turns to the required heading, and returns to the launch site oran alternate base through powered and gliding flight. At no timedoes the booster fly faster than M = 2.95; experiences Machnumbers greater than 2.0 for only 212 seconds. This. avoidance ofa hostile flight environment enables the booster to beconstructed of conventional materials without a thermalprotection system. This is illustrated in Figure 3-40.
3.7 Hypersonic Interceptor - Model 1074-0006
3.7.1 Concept Description
The vehicle, illustrated in Figure 3.41, has an overalllength of 169 ft 2 in and a wing span of 63 ft 4 in. The winghas a leading edge sweep of 720 on the inboard section and 500 onthe outboard section, a reference area of 2,085 ft 2, an aspectratio of 1.923, and a constant wing thickness ratio of 3.5%.
The airplane is designed for a one-man crew. Locatedforward, aft, and below the crew compartment areavionics/electronic equipment bays. Included in the 1,200 lbs ofavionics equipment are target acquisition, communication,navigation and identification, information management, anddefense functions. ECS equipment, oxygen, andelectrical/hydraulic subsystem equipment are located in the
105
I- C9Lua
I-I
a-4
LUL
40
L..
00
Ce 4
Lu .
16n6
.. N
i-
%, I
411
I WI
Lon
107
fuselage aft of the pilot. The body fuel is carried in integralinsulated tanks with a capacity of 53,000 lbs of liquid hydrogenfuel.
The vertical fin has an area of 240 ft 2 , a leading edgesweep of 600 and an aspect ratio of 0.98.
The horizontal tail has an area of 578 ft 2 , leading edgesweep of 60° and an aspect ratio of 1.75.
3.7.2 Aerodynamic
Estimated aerodynamic characteristics of the Model 1074-0006are presented in Figures 3-42 through 3-44.
Trimmed drag polars are shown for typical subsonic,supersonic and hypersonic flight conditions are shown in Figures3-42 through 3-44.
In Figures 3-42 through 3-44, the drag generated by PWSIM iscompared to LAAP (Large Airplane Analysis Program) and APAS(Aerodynamic Preliminary Analysis System) computer codes atsubsonic and supersonic speeds and to APAS at hypersonic speeds.
3.7.3 Weights
The weight statement for the 1074-0006 is shown in Figure 3-45. Weight estimation ground rules and assumptions are listedbelow:
"o The majority of aircraft structure is advanced hotstructures, capable of enduring the high temperatures ofsustained hypersonic flight
"o Airframe Integrated Nozzle and Inlet
"o Fly-by-wire surface controls
"o Avionics equipment as described in Section 3.7.1
"o Internal weapon carriage on two rotary launchers
"o Final aircraft geometry is the result of aerodynamic andweight parametric trade studies and represents the bestcompromise for overall performance
"o Judicious location of missiles and fuel such as tominimize CG travel as these items are expended
Uninstalled engine performance was computed using theGeneral Electric tandem turboramjet hyperjet, GE16/F40 study Bi.The engine is a low bypass ratio, hydrogen fueled augmentedturboramjet having a max augmented thrust of 57,718 lb sea levelstatic. The engine cycle characteristics are bypass ratio (BPR)= 1.5, overall pressure ratio (OPR) = 25, turbine inlettemperature (TIT) = STOICHIOMETRIC.
The inlet is located under the fuselage, centerline mounted.It is a two-dimensional, mixed compression inlet.
This inlet has a fixed first ramp, a flexible second ramp,and a movable third ramp. The boundary layer is controlled bymeans of porous bleed on the second and third ramp surfaces,sideplates, and a throat bleed slot located aft of the normalshock. The throat slot also acts as a bypass to remove excessinlet airflow for matching engine airflow demand with inletsupply and controls the position of the throat shock. The inletcapture area if 24.40 ft 2, sized for air requirements at Mach 5,100,000 feet.
The aftbody of the interceptor serves as the expansionsurface for the engine. Also, there is a turning vane which isused to maintain flow attachment of the exhaust plume on theaircraft aftbody throughout the aircraft flight regime.
3.7.5 Performance
The aircraft was configured to provide low drag at thedesign Mach number of 6.0. A design mission was specified (seeFigure 3-46) that involved flight at altitudes greater than100,000 feet, and sample results are shown in Figure 3-47.
4.0 Sample Results
This section contains an example PWSIM output (Figure 4-1).The output which is for the tactical fighter consists of:
o namelist inputs
o mission definitions
o airplane design (geometry) summary
o group weight statement
o weight design data and sensitivities
113
"o detailed weights
"o minimum profile drag
"o wave drag
"o drag due to lift
"o zero lift drag versus Mach number
"o mission results
"o level flight performance
"o engine data
"o inlet tables
"o aftbody drag tables
"o installed engine performance
"o airplane inlet maps
"o airplane afterbody maps
114
e '4.'
0
S06Go u.
0.w
001.
1.4
00
I CD:a 04
0. 'n 4~ A 04E- W M
>4 C"0> 0 14
W.I 0 1 4 Q l
rid~~6 (D~~C.) 0
V NU)nOD Mr- m (Y) N
I N Ir- 4o I I I M n I*!N NClOm 14v r-r-ra
000 v@0O6•00
• O0OO000r
N~- N4 N0>MDt
-4-4 cq woo a LCD
'IOV 1w
N HO0•f
E-4 %D OD 0 a %0%* %otfLn rincv) 14
C 0 400DOD H00
0
4.)
r- H0• ")M0 >C 000
••00
0- 0 - ,-I 05.C400
. . .. . . . . . .
w %0 % t 41o
, O % O I I H I00 i0 OD ODC r-4 >
02
04 W D% )IN N 04 Hr,%DN -4 a
N004C~4ChOhD ODODr- r,i-lw-r-f r-I
E-0 C>( 0 0 In W-4o au D C4 Cr)
0 0%0 C) ITOD " C) (r)
H o a o v qqC>1-
NI 0 40 N H -4 -1 C)
(Y)
04-i ) -H ) 0 - 0D:54-I Hx0-r F37e 0 :j0)0 - tto tou *r-i 0~~- M 4 )
0) CI)~~F E4-4O )0lH
H 0 H4(flrvu,)%or-coo) CzH
116
C.
vz
0 0
7 7h
Z 0 L LiU
C) I U. z
Z x j
X Z Z > 2
11
LUCL
o o-
0 0M. - 4-
31 L6 -ct lz w - It
w m 0- tvt, -
IL a z m j x0 2 . U.U. o a 4 Q 0 a39 3 N 0 w x x O a 4 w x I -
118,
IC-
I-L
* (d * W .0 0 0
PC. O0 4N m4 -4 -) - - - -n - - - - --
. . .. . .N . . . . . -
- 0 Z O. N 3 ~ UZZ2~Ow0 ý X U4 44424 0 6 w - ~ U i cz = U) C w
~ c 6. L6L 4 V L) X) N 0J AsiU U ~ U U - n i n i n in i n i
MOM&.0la -0 a M Vr-ONr-M#DV -41 ..... .. . . . IA n 'f*I0W0 40.3 I.-w nc..- - ----- -------------- mr' vm mm m.. J& 44D 40 - -- - -
145
00 *Nmmmac
8 0) pm um
C4 CV W to
. C) CD - -- 00 *OC r-t
0 0 0 N ,j .0r
I- 8§ 000WOw~ Pw w . . 0
08 8§ NIoN- .-w
4 0 0 NV 40I C1
Ii - ow0p 1 . l0
Wa Ij qj 8 0 0 5 w-cc0 43- In -. ,II..r - WO .
4~ ~ 14 4, 1 008W~ N WO 0 19 110. 8 .- cc 0) coCo0CL 04 ) Ow Qn -- N O )C
Or N .-9 01 0N 1-N4 ow AJIll 8 8 M~ W S 10 0 0 O05 W NLai ~ W O U 'N '-0 '.0mor ICW D 8 m -e) (4f. ry- l nt P.
N w C4 -
ILL 8 (n P C4= ID V- W4 0 O wcvinc4l -j -) ... w 0
- SW M0 0 w (0 aV10 CL x 0
z0 In 8 So10 W
U,~~ co 0) t-l 4A- 8l = N Z 0. OUZ - 4 oo In pW(- C - tn T - *oom Ox V . c- a .P 1W v M O w- *. Cno*g0 0 - 3w vI to 4 m
z 0 of- to Cj I w w4 z 0 - v-
00
0 w - ef 010 -(9 ovw- 0. Z O.01 V. r-WM O 8 §~ 6 0 '1=0 40) 4 BI1 a" Zm m .0In I-l .- . . LL. bS U
(9 S 0 - -to 0 0 8 v I- M: 1-" 0w 8
00 31 van u 101 C. (U in. - L U 04Z U CL u F-A .( 0
8 -N '- ( w no5 Li ot) 3'WWfl O 00U418 1 j wp x o 0 1' .7. 88 0 8m- Mi~ OCL N-U W- i U) I OM-
OW 0 i -j 4U 0 '4 .j F F06- 39 El '(. 0 - S4 00- 00 oui Z ms 31 4.- 00z ~&L I-u 9L U - IF- UN 0- 1w Z;0 U U0- 0.U 0. -1 0S - US W . U w u0.. -A US U L'- L. -4 0 a5 5 U)- u
0. 0 Z 0. Z Oia z -u cc 10 0 4O.00,UA-.0... 1j w 3 " - W - 44 CS W4 Wý 1 3 UU
om~~ ~ ~ 0U a 0 Wýc ob -
(0 " , 2 z iz0. 7 Z UAU146
.010 0.
80 00-0
LU 4w-jc
00z 8( 0
LU -
LU co
8 .0 (94n~ 0
000-.
z 0. In
S U. P- 8
80
Z 0. 0z mf u1
IX Ix L
I. LU 4
z Ix- .Z U I
40 00 9L0L
4147
8~~~0 00::::::8 ::::::: ten,:
N *444. 4..e C
*'r ..... to m &P . ci. v
oU *. *4Ci4. h Cr.. V
tn 4. . fl Inch In I4 C
0- .4.. 4 4 4 4
.4.. 4 C4 C4 r-4
I-uj o - 4-to w4 i0n in. in. 6h. win
9 co. *0 . * * * *
-
4-
C44
N 004
- 0 Z ID * r-. ;oq4 I IinnIgvti
w .V4: *IN v.o4 .0. tn in* C41 4 --nQ1eM
- 0 * . . Dan. anm -nmqwq
CI .4148
, * 4444444 -:. fl
* 4 .44 4. V 4... 4
t* on 4.. 4. 4.N . r.
oo C**4In,*440 U)o4 C4.4 .4 C4 4 C4N
'IC
inin 4.4. 4 4 4 * -
4 4o .4 .4 . . . 4 .
- - 4..~;C. ~ - .C
z n 44 4 4
InD *444'00(Din
CM4m44 C4 '
I.-S.-
W 149
0Wl 0
8 8,N C4 40
008 *m~o~m Uwtov~lf
0 0 r-w
00- 0Uomin W aco0tn~win N V) -W 9)o ~ c 2~ q0 w'0 U)~0-~ cc c.Wowrio-uO C
two w W-N-8 ; c4l
.0)DV 0000000 008 0000C4 o .. . -4 . .. 4. .
t0 08 -O~ 8 mo m-o r.-at' 0 M"M 8 womw -00--
. 0*Wv -00000000 . 000)0000
-~4 Gb0~0tflo COD 8l-,Om4 0 0Coz>;Ffl00 00000
aD~111 0 . .0 W z~ w :t N 0 NLo 0 011 ol n0 )V 0 c,0 W O mLn 0 r),WYI 0 0
0 ~~~ 00000-ooooo -000o0000-
CL 0 Q801toO- 0 nmfvw
W -wcl)t0n~fOO --. 0000 000; 0 0 000000 Q~ (N Im Z NW to..
C4 -4 _jf-wI w w w0§~ nlc u "8"O~n 8 . )c 0 C. r-; ~ mmW4o 4o 0 t 0 m~onO;0-cn I 0 08"m
OCntocn 0 000000000 . 00000000
4 iN u 0. *Oom 0000 0 ,mm000000 0 WC)0Nt -
N u z toIw& 0 0.m-m0 C~))) 0 CP W0
8§.00 0 a 0 LM cl-We~ 0 8 0 q~~4 Z (N ,w0 000000060800
U. -L)o-I
us 0 . 0 0 7 oo m m m o w o-w r cLn U 0. --------- 00 roo -owQ-------
N . .* N7'400 -0000000000 ;0000000000 ~ ~ t 00.* ~ i
- 0 Z ION .4 2 A ow 00z . . .. .. .W
C," 00IA 0. 0 0
0 3D 0 -z0 0 OwW0UU I- . uj 20U IZ, fN N N 0t00~~~ -= 0 0 . 00Uo ~a NW 0 .-0C 0~00-~ 1. §u am~
