^>TELcDYNS RYAN AERONAUTICAL - NASA · ^>telcdyns ryan aeronautical 2701 harbor drive, san diego....
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NASA CR-112323 N 7 3 24057 FEASIBILITY STUDY OF MODIFICATIONS TO BQM-34K DRONE FOR NASA RESEARCH APPLICATIONS ASTM 72-40 27 DECEMBER 1972 PREPARED UNDER CONTRACT NO. NAS1-11758 for NATIONAL AERONAUTICS AND SPACE ADMINISTRATION ^>TELcDYNS RYAN AERONAUTICAL 2701 HARBOR DRIVE, SAN DIEGO. CALIFORNIA 92112 AREA CODE 714/291-7311 https://ntrs.nasa.gov/search.jsp?R=19730015330 2018-08-06T16:09:35+00:00Z
Transcript of ^>TELcDYNS RYAN AERONAUTICAL - NASA · ^>telcdyns ryan aeronautical 2701 harbor drive, san diego....
NASA CR-112323
N 7 3 2 4 0 5 7
FEASIBILITY STUDY OF
MODIFICATIONS TO BQM-34K DRONE
FOR NASA RESEARCH APPLICATIONS
ASTM 72-40 27 DECEMBER 1972
PREPARED UNDER CONTRACT NO. NAS1-11758
for
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
^>TELcDYNS RYAN AERONAUTICAL2701 HARBOR DRIVE, SAN DIEGO. CALIFORNIA 92112 AREA CODE 714/291-7311
https://ntrs.nasa.gov/search.jsp?R=19730015330 2018-08-06T16:09:35+00:00Z
Prepared byH. A. JamesDesign Specialist
Approved byp. E. DiBartolaManagerPreliminary DesignAdvanced Systems
Approved byif. G.v TimmonsDirectorAdvanced Systems
ii.
FOREWORD
Contributions for this report on various subjects v/ere made by thefollowing personnel for Advanced Systems: D. C. Harper and II. M.Allen, design; D. .O. Nevingcr, \veights; R. W. Thompson, parametric,wing sizing; II. A. James, Project Leader, aerodynamics and perform-ance; F. L. Miller, systems;-!. Wilcox, reliability; and N. M. Bowers,structures and materials.
in
SUMMARY
The feasibility of modifying an existing supersonic drone, BQM-34E,into a NASA free-flight research vehicle is examined in this study. Thisremotely controlled vehicle would be capable of free-flight validationtesting of wing configurations representative of a wide range of researchapplications for advanced transports and fighters as well as RPVs. Thisstudy is addressed to three main topics per Contract No. NAS 1-11758,i.e.: aerodynamics and performance, design and structures, andcommand and control system.
Appropriate structural and control system modifications, reliability andoperational considerations, and ROM costs indicate that the BQM-34Edrone is indeed suitable as a NASA research vehicle.
During the initial portipn of the study, wing sizing to specified aerody-namic and performance criteria was accomplished. This resulted inthe definition of six point designs matched to the modified BQM-34E withits basic propulsion system. From these results, NASA selected arepresentative research configuration for more in-depth structuraldesign and control system studies.
The structural design studies identified several alternative engineeringsolutions for the testing of high and low-wing configurations. These wereevaluated in terms of cost, complexity, and model similarity. Repre-sentative control and high-lift devices were configured for transonicflutter mode suppression research testing. Practical methods ofachieving variations of wing bending and torsional rigidity were identified.
The results of a comprehensive anal3'sis of command and control systemsrequired for various types of research programs are summarized. Thebasic control and AFSC system is amenable .to modifications with existinghardware to accommodate steady-state as well as dynamic loads, flutter,and variable-stability research programs.
C
CONTENTS
SECTION PAGE
1.0 INTRODUCTION 1/
.2.0 TECHNICAL APPROACH 3
3.0 RESULTS ' . 5
3.1 Preliminary Studies 53.2 Point Designs 213.3 Research Configuration 643.4 Support Studies - 169
4.0 ' CONCLUSIONS . 183
5.0 RECOMMENDATIONS ... . 185
6.0 NOTATIONS AND SYMBOLS 187
7.0 REFERENCES 191
vil
LIST OF FJG U'R ES
FIGURE PAGE
3-1 AVSYN Results 63-2 AVSYN Results, Sea Level Launch / 73-3 AVSYN Wing Parameter Study /: ' 83-4 NASA Wing Study, Preliminary Estimate of the
Thrust Required and Available for the No. 1 Wing-Design 9
3-5 NASA Wing Study, Sizing Study for Wing No. 1 103-6 NASA Wing Study, Preliminary Estimate of the
Thrust Required and Available for the No. 2 Wing-Design _ _ 11
3-7 NASA Wing Study, Sizing Study for Wing No. 2 123-8 NASA Wing Study, Preliminary Estimate of Thrust
Required and Available for the No. 3 Wing Design 133-9 NASA Wing-Study, Sizing Study for Wing No. 3 143-10 NASA Wing Study, Preliminary Estimate of Thrust
Required and Available for the No, 4 Wing Design 153-11 NASA Wing Study, Sizing Study 'for Wing No. 4 16
(._ 3-12 NASA Wing Study, Preliminary Estimate of ThrustRequired and Available for No. 5 Wing Design 17
3-13 NASA Wing Study, Sizing Study for Wing No. 5 183-14 NASA Wing Study, Preliminary Estimate of Thrust
Required and Available for the No. 6 Wing Design 193-15 NASA Wing Study, Sizing Study for Wing No. 6 203-16 No. 1-30 Mach Number vs. Altitude 253-17 . 'No. 2-30 Mach Numbei vs. Altitude 263-18 No. 3-24 Mach Number vs. Altitude 273-19 No. 4-40 Mach Number vs. Altitude 28
" 3-20 No. 5-60 Mach Number vs. Altitude 293-21 No. 6-35 Mach Number vs. Altitude . 303-22 No. 1-30 Specific Endurance 313-23 No. 2-30 Specific Endurance 32
. 3-24 No. 3-24 Specific Endurance 333-25 No. 4-40 Specific Endurance 343-26 No. 5-60 Specific Endurance 353-27 No. 6-35 Specific Endurance 363-28 No. 1-30 CDO vs. Mach Number . 47
IX
LIST OF FIGURES (Continued)
FIGURE ' PAGE
3-29 No. 1-30 Induced Drag Coefficient vs. Mach Number 433-30 No. 1-30 Longitudinal Characteristics 493-31 No. 2-30 CDO vs. Mach Number 503-32 No. 2-30 Induced Drag Coefficient vs. Mach Number 513-33 No. 2-30 Longitudinal Characteristics 523-34 No. 3-24 CDO
vs- Mach Number ' 533-35 No. 3-24 Induced Drag Coefficient vs. Mach Number 543-36 No. 3-24 Longitudinal Characteristics 553-37 No. 4-40 CDO
vs- Mach Number 563-38 No. 4-40 Induced Drag Coefficient vs. Mach Number 573-39 No. 4-40 Longitudinal Characteristics 583-40 No. 5-GO CDO
vs- Mach Number 593-41 No. 5-60 Induced Drag Coefficient vs. Mach Number 603-42 No. 5-60 Longitudinal Characteristics Gl3-43' No. 6-35 CDO
vs- Mach Number 623-44 No. 6-35 Longitudinal Characteristics 633-45 General Arrangement, 1-30-2 673-46 Design Alternatives .683-47 Wing/Fuselage Configuration Tradeoff 693-48 Wing Installation and Fuselage Modification 1-30-2 713-49 ' Wing Installation and Fuselage Modification 1-30-1 723-50 Inboard Profile, Configurations 1-30-1 and 1-30-2 753-51 Area Rule Modifications 773-52 Area Distribution for the Basic BQM-34E 783-53 Area Distribution, Configuration 1-30-2 NASA
Research Wing, Model 166 793-54 Wing Plan Form Structural Arrangement 803-55 Wing Structural Cross Sections .. 813-56 Mach Number vs. Altitude - 833-57 V-n Diagrams - Symmetrical Maneuvers, Model
BQM-34E . 843-58 V-n Diagrams - Unsymmetrical Maneuvers, Model
BQM-34E . . 8 53-59 Model AQM-34R Drone 1243-60 Analysis Approach and Cycle 1253-61 Structural Analysis Procedure 1273-62 Concept 1, .Variable Wing Stiffness 1313-63 Concept 2, Variable Wing Stiffness 132
. 3-64 Concept 3, Variable Wing Stiffness 133
LIST OF FIGURES (Continued)
FIGURE ° PAGE•
3-05 Concept 4, Variable Wing .stiffness 1343-GG Concept 5, Variable Win;; S t i f fnes s 135
., . 3-67 Concept G, Variable Win;.; St i f fness / .1363-68 Concept 7, Variable Wing S t i f fness ' 1373-G9 Wing Stiffness Variation, Model 147TF Composite
Skins . 1 3 93-70 Deleted3-71 Estimated Static Stability and CG Range 1433-72 Model BQM-3-1E With Supercritical Wing 1-30,
Elevator Angle Required for Trim 145.' 3-73 Model BQM-34E With Supercritical Wing 1-30,
Estimated Maneuver Capabili ty 1463-74 ' Model BQM-.'ME Est imated Change in Stability Due
to Added Fuselage Section Forward of Wing 1503-75 ModelBQM-34E With Supercritical Wing, -
Comparison of Lateral-Directional StabilityDerivatives 151
/"' 3-76 No. 1-30-2 Time History ojt a Typical NASAResearch Mission 153
3-77 No. 1-30 Maximum Rate-of-Climb vs. Altitude 1543-78 No. 1-30 Specific Range vs. Mach Number, 2000
Pounds 1553-79 Avionics Functional Bloc!; Diagram 1573-80 Guidance and Control Concepts 1603-81 Longitudinal Axis AFCS Channel 1643-82 Lateral Axis AFCS Channel " 1653-83 Model BQM-34E Reliabi l i ty Functional Block
Diagram, Operational Complex 1713-84 NASA Research Vehicle Recovery Phase Reliability,
Functional Block Diagram " 1735-1 Recommendations for Phase II "186
LIST OF TABLES
TABLE PAGE
1-1 Wings Planned for Study / 2 ;3-1 NASA Research Mission Tabulation, Wing
No. 1-30, Mach 0.98 Transport 383-2 NASA Research Mission Tabulation, Wing
No. 2-30, Mach 0. DO Transport 383-3 NASA Research Mission Tabulation, Wing
No. 3-24, Air-to-Air RPV 39. 3-4 NASA Research Mission Tabulation, Wing j
No. 4-40, Endurance Turbojet 39 i3-5 . . NASA Research Mission Tabulation, Wing i
No. 5-60, Endurance Turbofan 40 I3-G NASA Research Mission Tabulation, Wing |_
\ No. 6-35, SST Configuration - - 403-7 Profile Drag Buildup, Model 1-30 413-8 Profile Drag Buildup, Model 2-30 42 i3-9 Profile Drag Buildup,-Model No. 3-24 43 . ' , '3-10 Profile Drag Buildup, Model No. 4-40 44 j3-11 Profile Drag Buildup, Model No. 5 453-12 Profile Drag Buildup, Model No. 6-35 463-13 . NASA Point Design Summary 653-14 ' Configuration List 663-15 Structural Design Criteria Summary, Parachute
Recovery .- 863-16 Structural Design Criteria Summary, MARS
Recovery 913-17 • Figures of Merit, Associated Parameters,
Weighting Factors 943-18 Comparison Matrix , 953-19 Evaluation 993-20 Weight and Balance Summary, NASA Wing Study
(Low Wing) 1213-21 Command Guidance Equipments 1623-22 Secondary Power 1683-23 ' NASA BQM-34E and Navy BQM-34E Flight
Phase Reliability Prediction Comparison 176
LIST OF TABLES (Continued)
TABLE PAGE
3-24 NASA BQM-34E and Navy BQM-3-1E RecoveryPhase Reliability Prediction Comparison 177
3-25 Estimate of Tasks 1793-2G Possible Elements of Flight Assurance/
Determination 181
xiv
1.0 INTRODUCTION-
is conducting intensive laboratory and flight test programs toenhance the development of both advanced civilian and military aircraft.In support of these programs, a relatively low-cost, remotely controlled,research vehicle could provide critically needed test data in a mostexpeditious manner and without the risk of human l ife. It would be parti-cularly valuable in the critical test and development phase, prior to theavailability of full-scale, manned research aircraft and/or, during thevalidation phase in support of corrections with wind tunnel test data. Insupport of these objectives, the purpose of this study is to determine thefeasibility of adapting the supersonic BQM-34E drone to accommodatea proposed free-flight research program which would include wings havinga broad range of applications.
At contract go-ahead, study guidelines and objectives were established ata joint NASA/Teleclyne Ryan meeting, as summarized in Reference 1.The proposed NASA research drone would be capable of accepting researchwings with a. broad range of subsonic and supersonic application.
This study encompasses only the conceptual.phase and defines, in general,the engineering approach and rough order of magnitude of resourcesrequired to modify an existing drone into a remotely controlled researchvehicle. The particular vehicle is unique in that it offers continuouslypowered flight test performance capabilities throughout the subsonic andsupersonic flight regimes, with reasonable endurance. From an aerody-namic standpoint, this configuration is representative of an ideal limit,in.terms of aerodynamic cleanliness.
In terms of structural integrity, this vehicle has a rugged airframedesigned to operate up to ultimate dynamic pressure of 2133 psf, whichcompares quite favorably with that of any of the known advanced fighters.
The current shoulder wing crossover structure is readily adaptible tointerchangeable wings at low cost. Low-wing installations are also feasi-ble at increased cost and complexity, depending upon emphasis .inaccordance with research priorities.
Preliminary design guidelines for sizing six possible research wings,based on a modified BQM-34E system, are summarized in Table 1-1.
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2. 0 TECHNICAL A PPROACH
The technical approach utilized in the initial portion of the study was toconduct preliminary design studies of wing configurations having a widerange of subsonic and supersonic applications. This was accomplishedfor six types of wings, in accordance with design and model similaritycriteria.summarized in Table 1-1. It will be noted that wings applicableto advanced subsonic transports incorporating supercritical wing tech-nology, an advanced-maneuverability fighter, an SST, and RPV areincluded.
The initial study guidelines included the following considerations:
a. Revision to internal fuel system for a.capacity of about 400pounds fuel.
•;
b. MARS or parachute recovery.
c. Conventional ailerons plus stabilizer tail.
d. . Air launch primary, ground launch secondary.
e. High and low-wing test capability.
f. Remote or onboard command and guidance systems.
g. Unique, one-of-a-kind, research vehicle.
In this portion of the study, tail volume coefficients and wing-body geometricsimilarity constraints were kept close to those typical for each wingappli- , ,,cation. This portion of the study was then summarized into a summarydocument designated as ASTM 72-22. This was submitted to NASA, alongwith three-view layouts and area distributions of each point design. Thisportion of the study provided NASA with a basis for selection of a configu-ration (1-30) for Tasks II and III. Engineering design and structuralstudies were then carried out for a feasible, one-of-a-kind, research drone,capable of testing a variety of high and low-wing configurations. Theassociated studies involving advanced structural materials, proportionalcontrol, and control law system capabilities were carried out on the basis
of the low-wing sonic transport wing configuration identified in this studyas wing 1-30-2. This study w:is concluded with .ROM costs and recom-mendations for an immediate follow-on program.
3.0 RESULTS
' The results of this feasibility study arc presented in Paragraphs 3.1through 3.4. These results are presented in a logical sequence, startingwith the preliminary parametrics and wing sizing studies and followed by.more in-depth engineering studies accomplished on a NASA-selectedrepresentative research drone configuration.
3.1 PRELIM1NARY STUDIES
Preliminary vehicle si/ing data \vas first explored by means of theTeledyne Ryan Advanced Systems vehicle-sizing program designated asAVSYN. The computerized program, AVSYN, can accept up to 145 designand mission variables to size remotely piloted vehicles quickly. Thefeasibility of accomplishing designated 20 to 31-minute missions with aNASA payloacl of 250 pounds with a reasonably sized vehicle based on theBQM-34E propulsion system was examined. Trends versus wing aspectratio and wing area are shown in Figures 3-1, 3-2, and 3-3.
The significant results from this portion of the study indicated the feasi-bilit}' of vehicle gross weights from 2500 to 3000 pounds and fuel loads ofabout 400. pounds, sufficient to accomplish the NASA mission requirements.
Wing-Sizing Study
The preliminary wing design criteria for wing aspect ratios, Mach num-bers, and l if t -coefficients "iYbm Table 3-1 were utilized to determine wingarea and Reynolds number versus altitude for an assumed fixed weight of2500 pounds.
.. The results of this sizing study are illustrated for each of the wing appli-cations in Figures 3-4 through 3-15. These data provided a range of
.<. wing areas to be considered for each of the applications. It was notedthat small wings were bounded by geometric body width to span constraintWb/b. At 1-g f l igh t condi t ion, results show that small wings achievedhigher Reynolds numbers than did the larger wings. An additional con-straint to provide longitudinal t r i m and stability involved horizontal appli-cable tail volume coefficients for each type of wing considered in thisstudy. Coordination with NASA (Mr. Ferris) confirmed our views that
Figure .'3-1. AVSYX Results
Figure 3-2. AVSYN Results , Sea Level Launch
Figure 3-3. AVSYN Wing Parameter Study
Figure 3-4. NASA Wing Study, Pre l iminary Est imate of the Thrust Requiredand Available for the No. 1 Wing Design
LU.,- W:yi.T\:^:Ji-'.-..-.
. ; . : ; • : ' : : : ; ) : • : I.:: \ A - . \ . \ \ - . ::.:..'..:.: :.:::!.I..L ..:/..• .L:.::..>;.. .:..„!.....:.
• ! • : : : - - - ; : - : : ! : - : / : f :; i;:::\ •
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vFigure 3-5. NASA \Vinij Study, .S ix ing Study for Wing No. 1
10'
Figure 3-G. NASA Wing Study, Pre l iminary Est imate of.the Thrust Requiredand Avai lable for the No. 2 Wing Design
11.
Figure 3-7. NASA Wing Study, Sizing Study for Wing No. 2
12
CFigure 3-8. NASA Wing Study, Prel iminary Estimate of Thrust Required
and Available for the No. 3 Wing Design
Figure 3-9. NASA Wing Study,. Sizing Study for Wing No. 3
14
Figure 3-10. NASA Wing Study, Pro l iminary Estimate of Thrust Requiredand Available for the No. 4 Wing Design
15
Figure 3-11. NASA Wing Study, Sizing Study for Wing No.4
16
X.TKAFOLATKD1ST THiUJ^r!
i
v.Figure 3-12. NASA Wing Study, Preliminary Estimate of Thrust Required
and Available for No. 5 Wing Design
17
Figure 3-13. NASA Wing Study, Sizing Study for Wing No.5
Figure 3-14. NASA W i n g Study, P re l imina ry Estimate of Thrust Requiredand Available for the No. G Wing Design
Figure 3-13. NASA Wing Study, S iz ing Study for Wing No. 6
horizontal tail volume coefficients for most applications should be atleast O . G O to 0.80. This criterion would limit most wing sizes to lessthan 45 square feet. In only one appl ica t ion , involving wing 5 (\vhich hadlaminar airfoils) was a deviation on tail volume coefficient permitteddown to 0.40, close to that of a similar vehicle in existence. An addi-tional flight limitation \vas the thrust-limited maximum altitude at thedesignated design Mach number and lif t coefficient.
The results of this portion of the study were then examined for compati-bility with the BQM-341-; fuselage crossover, center of gravity, and tailarms by means of three-rview design layouts. The design layouts includedthe following range of feasible wing areas for each application:
WING NO. WING AREAS, SQUARE FEET
1 30 to 502 2G to 503 20 to 284 40 to 605 ' 40 to 606 25 to .35 '
3.2 POINT DESIGNS
The preliminary design guidelines for developing feasible designs foreach of the six applications consisted of the following:
a. Wing crossover structure close to that of the basic vehicle.
b. Wing c/4 close to Station No. 264 to achieve reasonablecenter-of-gravity balance.
c. Horizontal volume coefficient, Vj-j s 0. 6 to 0. 8.
d. Vertical volume coefficient Vy a 0.08.
e. Revision of fuel system to about 400-pound fuel capacity.
f. Air launch primary, ground launch secondary.
g. Conventional aluminum riveted construction or equivalent.
h: Conventional ailerons plus stabilizer tail.
i. MARS or standard parachute recovery secondary.
21
It was apparent at-thc onset of this study that a high-wing configurat ioncould more easily be developed than could :i low-wing configuration.However, it was considered desirable l<> achieve a low-wing capabili ty,since tliis would be more representative of transport configurations.
/X
WEIGHTS ANALYSIS
The weight of the basic BQM-34E, less wing and target augmentat ionequipment, is tabulated below:
•ITEMWEIGHT(pounds)
Wing GroupTail GroupBody GroupTakeoff and Recovery EquipmentPropulsion .Lube and Fuel SystemElectricalControlsGuidanceElectronics
.Environmental Protection Equipment
50.0273.3122.0427.0
36.1139.4
36.742.8-50.6.10.1
Weight Empty - RevisedUnusable Fuel and OilRefrigerant System
Zero Fuel Weight - RevisedInternal FuelRefrigerant
Gross Weight - Revised
(1188.0)15.220.6
(1223.8)274.0
8.3
(1506.1)
Basic Items RemovedWingTarget Augmentation
142.2171.8
22
Modifications Weight Summary
Estimated weight for ant icipated modifications to the BQM-34E arctabulated below:
/
WEIGHTITEM (pounds)
NASA Payload 250.0Two Span Ailerons 20.0Additional Fuel 76.0Additional Tani<age 24.0High-Li ft Devices 50.0MARS Recovery System 50.0Wing Crossover Adapter 50.0Revised Air Launch Fittings 10.0Area Rule Modifications 50.0Ballast Provisions 100.0
Total Modification Weight (680.0)
Estimated Modified Vehicle Gross Weight
The estimated vehicle gross weights for each wing configuration aretabulated below:
. ' WEIGHTCONFIGURATION (pounds)
Configuration 1 (Sw = 30 ft. 2)
BQM-34E G\V Revised 1506.1Modifications 680.0Wing. 156.0
Gross Weight (2340.1)
Configuration 2 (Sw = 30 ft.2)
BQM-34E GW Revised 1506.1Modifications 680.0Wing 159.0
Gross Weight (2345.1)
23
CONFIGURATION
Configuration 3 (Sv/ = 24 ft.2)
BQM-34E GW RevisedModificationsWing
Gross Weight
• Configuration 4 (Sw = 40 ft .2)
BQM-34E GW RevisedModificationsWing-
Gross Weight
Configuration 5 (Sw .= GO ft .2)
BQM-34E GW RevisedModificationsWing-
Gross Weight
Configuration 6 (Sw = 35 ft. 2)
BQM-34E GW RevisedModificationsWing
Gross Weight
WEIGHT(pounds)
150G.1G80.0128.0
(2314.1)
1506,1680.0199.0
.(2.385.1)
150G.1680.0226.0
(2412.1)
1506.1680.0136.0
(2322.1)
Performance" Envelopes
The general performance and typical NASA research mission capabilitiesof each wing application developed from the design study were examinedin this portion of the study (Figures 3-16 through 3-27).
NOTE
The notation for each design includes a wing numbercorresponding to its application in Table 1-1. Thedash number denotes \slng area; i.e., 1-30 is wing 1with a 30-square-foot wing.
24
Figure 3-16. No. 1-30 Mach Number vs. Altitude
25
Figure 3-17. No. 2-30 Mach Number vs. Altitude
26
Figure 3-18. No. 3-24 Mach Number vs. Al t i tude
27
Figure 3-19. No. -1-40 Mach Number vs. .Altitude
28
Figure 3-20. No. 5-60 Much Number vs.- Al t i tude
29
Figure 3-21. No, G-35 Much Number vs. Altitude
30
Figure 3-22. No. 1-30 Specific Endurance
Figure 3-23. No. 2-30 Spec.ific Endurance'
32
Figure 3-24. No. 3-2-1 Specific Endurance
33
.;T-r. .;
Figure 3-25. No. 4-40 Specific Endurance
34
Figure 3-2G. No. 5-GO Specif ic Endurance
Figure 3-27. No. G-35 Speci f ic Endurance
36
The performance evaluation was preceded by estimates of the requiredlongitudinal aerodynamic coeff ic ients versus Mach number and angle ofattack. Avai lable methods included in the AAF Datcom Handbook, NASAReports, and Tcledyne Ryan estimation methods were applied directly asnecessary to generate aerodynamic coefficients for this study. For mostof the subject configurations, wind tunnel test data , due to compressibilityand flow separation phenomena, were available as the data basis.
The results of the flight envelope capabilities evaluation of each pointdesign, at three typical weights, are included in Figures 3-1G through3-26. Typical. NASA research mission capabilities are presented inTables 3-1 through 3-6. Examples of the subsonic drag bui ldup for eachof the wing conf igura t ions are included in Tables 3-7 through 3-12 atdesign altitudes. The supersonic wave drag and subsonic drag divergencephenomena were estimated by available Tcledyne Eyan Advanced Systemsempirical methods. , ,
The applicable longitudinal coeff ic ients versus Mach number for each ofthe subject configurations u t i l i zed to determine the flight performanceenvelopes are included in Figures 3-28 through 3-44.
The flight performance capabilities of each configuration \vere determinedby means of the computerized Teledyne Eyan performance programs.
Area Rule Modifications
The distribution of volume in terms of cross-sectional area versuslength was identif ied for each of the six-point designs. It was noted thatthe equivalent body fineness ratio of each point design was quite high, sothat even without ideal area rule modifications this research vehiclewould be expected to have low wave drag, i.e.:
CONFIGURATION EQUIVALENT BODY F.R.
1-30 . 13.02-30 12.63-24 13.94-40 12.25-60 10.06-35 13.8
The variation of cross-sectional area of the selected configuration 1-30-2can be compared with a recommended and an ideal zero-lift distributionin Figures 3-52 and 3-53.
37
TABLE 3-1
C
NASA R E S E A R C H MISSION TABULATION,WING NO. 1-P.O, MAC11 0.98 TRANSPORT
MISSIONSEGMENT
12
3
SEGMENT DESCRIPTION
Warm up and launch at 10, 000 fl.Max. climb to 45,000 ft. at Mach 0.9Design cruise at Mach 0.9S at 45,000ft .
t(min.)
/ 0.02 .2
25.8
WEIGHT(lb.)
45.053. 1
246.9
R
(nm)
0.018.0
242 . 0
Launch Weight:Fuel Weight:Zero Fuel Weight:
2342.1 lb.350.0 lb.
1992.1 lb.
TABLE 3-2
NASA RESEARCH MISSION TABULATION,WING NO. 2-30, MACH 0.90 TRANSPORT
MISSIONSEGMENT
123
SEGMENT DESCRIPTION
Warmup and launch at 10, 000 ft.Max. climb to 50, 000 ft. at Mach 0,9Design cruise at 50, 000 ft. at Mach0.9
t(min.)
0.03.83
49.9
. WEIGHT( lb.)
45.090.4
214.6
R(nm)
0.032.9
429 . 0
Launch Weight:Fuel Weight:Zero Fuel Weight:
.2345.1 lb.350.0 lb.
1995.1 lb.
38
TABLE 3-3
NASA R E S E A R C H MISSION TABULATION,WING NO. 3-24, AIR-TO-AIK HPV
MISSIONSEGMENT
123
SEGMENT DESCRIPTION
Warmup and launch at 10,000 ft.Max. climb to' 40, 000 ft.Design cruise at Mach 1.4 at 40, 000ft .
t(min. )
/ O . O2.55
12.38
WEIGHT
(Ib.)
45.070.8
234.2
R '
(nm)
0.021.9
165.6
NOTE: Add 50.5 pounds of fuel to increase secernent No. 3 to 15 minutes.
Launch Weight:Fuel Weight:Zero Fuel Weight:
2314.1 Ib.350.0 Ib.
1964.1 Ib.
BLE 3-4
NASA RESEARCH MISSION TABULATION,WING NO. 4-40, ENDURANCE TURBOJET
MISSIONSEGMENT
123
SEGME NT DESCRIPTION
Warmup and launch r t 10, 000 ft.Max. climb' to 50, 000 ft.Design cruise at Mach 0.9 at 50, 000ft. -
t(min. )
0.04.54
47.3
WEIGHT(Ib.)
45.0- 99.7
205.3
R .(nm)
0.043.4
406.0
Launch Weight:Fuel Weight:
23S5.1 Ib.350.0 Ib.
Zero Fuel Weight: 2035. 1 Ib.
39
TABLE 3-5
NASA RESEARCH MISSION TABULATION,WING NO. 5-60, ENDURANCE TURBOFAN
MISSIONSEGMENT
123
SEGMENT DESCRIPTION .
Warm. up and launch al 10,000 ft.Max. c l imb to 55,000 ft.Design cruise at 55,000 ft., at Mach 0.75
t( in in . )
/ 0.05 . 13
58.0
WEIGHT( Ib . )
45.0105.9199.1
R(mn)
0.034.3
418.11
Launch Weight: 2412.1 Ib.Fuel Weight: 350.0 Ib.Zero Fuel Weight: 2 0 G 2 . 1 Ib.
TABLE 3-6
NASA RESEARCH MISSION TABULATION.,WING-NO, G-35, SST CONFIGURATION
MISSIONSEGMENT
1
23
SEGMENT DESCRIPTION
Warm up launch at 10,000 f t . -Max. climb to 40,000 ft.Design cruise at Mach 1. 4 at 40, 000ft-.
t(min.)
0.0' 2.84
12.54
WEIGHT( Ib . )
45.078.7
2 2 G . 3
R(nm)
0.026. 2
167,5
NOTE: Add 44.3 pounds fuel to achieve 15-minute segment No. 3.*
Launch Weight: 2322. l ib .Fuel Weight: 350.0 Ib.Zero Fuel Weight: 1972. l i b .
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cTo achieve an ideal distribution of volume, the fine-body cross-sectionalarea would have to be almost doubled. A vei*y 1'easonablc compromise,well suited to the Maeli range of this research drone and designated as"minimum area rule fair ings", is illustrated in Figure 3-51. Theshoulder fairings provide the least interference with access doors andlaunch and recovery f i t t ings . In any case, this docs not appear to be aserious consideration for this particular configuration, due to its inher-ently high equivalent both' fineness ratio.
3.3 R ESEAHCII CON FIGURATION
The results of the wing parametric and sizing analyses summarized inTable 3-13 were included in an interim report to NASA (ASTM 72-22).From this data, a representative research drone configuration (Figure3-45) was selected for more in-depth engineering studies, to be pre-sented in the ensuing sections. The results of structural and designstudies, as well as pertinent features of a command and guidance systemcapable of accomplishing a variety of NASA research tasks, is includedin Paragraphs 3.3.1 through 3.3.7. Various tests required to achieveassurance of success are summarized in Paragraph 3.4.
3.3.1 Wing Location
Four wing location concepts were investigated, all with the same basicwing planform (configuration 1, 30 square feet) which was selected from:the previous six-wing parametric study. These four were identified asconfiguration 1-30-1, 1-30-2, 1-30-3, and 1-30-4 (Figures 3-46 and3-47), as shown in Table 3-14.
A tradeoff study was performed to select the best all-around method ofwing installation (Figure 3-47). Main7 parameters were considered, butthey were reduced to four significant ones: transport geometric.simula-tion, vehicle performance, estimated cost, and operational factors.Transport simulation consisted of determining how closely the design ofthe drone could be scaled to be representative of the proposed supercriti-cal wing transports. Flight duration, stability, drag, etc., were con-sidered under the vehicle performance aspect. In the cost estimate,configurations were considered in the light of Teledyne Ryan's experiencewith them, their total costs, and other factors. Air launch and recoverydifficulties, as well as the chance of damage in case of ground impact,were considered in the operational aspect of the study. The results(Figure 3-47) indicate that configuration 1-30-2, the low midwing, is thebest, with configuration 1-30-1 the second best. Since both of theseconfigurations were of interest to NASA,- they were both investigated as
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.to the method of construction. In addition, it was determined from NASAthat tlie ability to convert one vehicle from a low wing to a high wing(configuration 1-30-2 to 1-30-1) would be desirable from a researchviewpoint. This capability was also included in the investigation.
The stretched fuselage configuration 1-30-3, with a low wing, was dis-carded because of cost, air launch difficulty, centcr-of-gravity travel,and some increase in wetted area. Its only re^l advantage was increasedfuel volume for longer flight duration. Configuration 1-30-4 was dis-carded because of increased drag due to an increase in the frontal areaand a greater chance of wing damage in case of ground impact.
Low-Wing Attachment (Configuration 1-30-2)
This configuration requires modifications to the existing BQM-34E in thearea of the fuselage fuel tank in order to-mount the wing panels externally.The modification will consist of removing two sheet metal frames at XF250. bGO and XF 258.340. Three heavier machined frames located atXF 247.800, XF 254.000, and XF 260.200 (Figure 3-48) will replace theframes. The replacement frames will be within the existing tank skinline, thereby avoiding any fuel tank sealing problems. A suitable materialthickness will be built into the frames to withstand the loads imposed fromwing bending and torsion as well as the heavy bosses required for concen-trated bolt locations. The wing panel will be tension bolted to theseframes through adapter fittings located outside the tank skin. Thismachined adapter (Figure 3-48, Section B~B) will be bolted to the inter-nal frames at six places, two in each frame, with 7/16-inch-diameter,300,000-psi, heat-treated bolts. To mount the wings to the adapter fitting,5/16-inch-diametcr, 300,000-psi, heat-treated bolts are used. Thedesired wing incidence and dihedral will be incorporated into the adapterfitting. Minor incidence and dihedral changes can be made by machiningalternate adapter fittings as required,
High Wing (Configuration 1-30-1)
This configuration can be considered as an alternate to the low-wingconfiguration as the primary vehicle., As an alternate to configuration1-30-2, the high-wing location can be created by removing the wings andadapter fittings from configuration 1-30-2 and installing a new, machined,wing carrythrough box (Figure 3-49). This torque box will fit betweenthe existing fuel tank top and the parachute riser trough. It will bolt intothe top of the machined frames added for configuration 1-30-2. Twelvebarrel nut and wing attach studs are furnished to allow interchangeable
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installation of each wing panel. Lightening holes are provided in thetorque box to allow wire and plumbing passage and a reduction in weight.
Fuel and Pay load Provisions
Configuration 1-P.O-l, High Wing (Figure 3-43). - The fuel for this con-figuration is stored in (he regular fuselage tank, with optional fuel avail-able in the nose and a small tank behind the wing carrythrough box.Assuming JP-5 fuel at approximately 6.8 pounds per gallon, the fuelweight for each tank is as follows:
Main tank in fuselage 263 pounds
Auxiliary in nose 70 pounds
Auxiliary,behind wing box 20 pounds
Total 353 pounds
The pressurized pay load provisions are all in the nose equipment com-partment, and available space exists around the essential equipmentlocated in the compartment. Available and optional volume locations andsizes may be summarized as follows:
Nose cone (available) 0.30 cu. ft.
Cooling system removed (optional) ' 0.75 cu. ft.
Shaker unit removed (optional) 0.60 cu. ft.
103-gallon fuel tank removed (optional) 1.40 cu. ft.
Miscellaneous volumes (available) l . lOcu. ft,
Total 4.15 cu, ft.
In addition to the nose compartment volume, there is approximately 0.5cubic foot available around the wing carrythrough box if the optional fueltank is removed. This area is unpressurized.
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Configuration 1-30-2, Low Wing (Figure 3-48). - The fuel and payloadprovisions are essentially the same as those of configuration 1-30-1,with the exception of the space above the fuel tank. Since no wing carry-
* through box is required, the fuel load can be increased by approximately20 pounds in the auxiliary tank over the main tank, thereby providing atotal fuel capacity of 373 pounds. In the case of payload provision, if
C
73
the fuel tank above the main tank is removed the available unprcssurizedpayload volume is approximately 1.0 cubic foot. A detailed inboardprofile of this arrangement is illustrated in Figure 3-50.
Area Rule Modifications
Vehicle area ruling can be accomplished by installing external fairingson the fuselage/wing joints. The fairings mil be fiber-glass structures(Figure 3-51) held to the fuselage with screws. Points requiring frequentservice may require special access doors through the fairings. Typicalvehicle area distributions are "shown in Figures 3-52 and 3-53, and vari-ations of these can be achieved by redesigned fairings.
3.3.2 Wing Design and Construction
Structural Arrangement
The wing (Figure 3-54) is a tapered, swept wing with a supercriticalairfoil section, 11 percent thick at the root and 7 percent at the tip. Thesweep angle is"40 degrees 21 minutes at the wing quarter chord. Inboardand outboard ailerons are utilized, along with a hinged leading edge cap-able of moving up and down.
The primary bending and torsional structure of the wing is composed ofa full-depth honeycomb box bounded by sheet metal channel spars at 15and 60 percent chord. A machined aluminum root fitting is located atthe root of the bending box, allowing wing removal from the mating fuse-lage fitting. The trailing-edge fixed structure is of fiber glass and isremovable for control system access. A sheet metal, removable, leadingedge is furnished for access to the movable leading-edge controls. Thewingtip is a foam-filled, fiber-glass shell bolted to the torque box close-out rib.
Variable-Stiff ness Wings
The honeycomb torque box wing construction (Figure 3-55) will permitwings of varying stiffness to be designed and built without any majortooling change. This type of wing has nearly all of the bending and tor-sional material concentrated in the upper and lower skins, which arebonded to the full-depth aluminum honeycomb core. By altering themodulus of elasticity, thickness, and (in the case of fiber materials) thefiber orientation from the \\ing elastic axes, the properties of these skinscan be varied considerably. Computer programs such as SQ5, LAP*,
74
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CONFIGURATION 1-30-1
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.1.4 10 FT
CONFIGURATION 1-30-2
Figure ?-50. Inboard Profile, Configurations 1-30-1 and 1-30-2
Page Intentionally Left Blank
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box beam, NASTRAN, and WAVES I are available and have beenused to aid in Die design of this type of wing. -Candidate skin materials,each having certain structural features, are glass fiber, magnesium,aluminum, PRD-49 fiber, boron fiber, steel, and graphite fiber. Thematerials used for skins outside the bending torque box need not change.The honeycomb and channel spars will likely remain aluminum, althoughother materials may be considered, depending upon wing requirements.//3.3.2.1 Preliminary Structural Design Criteria
The basis for the preliminary structural design criteria for the modifiedBQM-34E research vehicle with the NASA supercritical wing shall be thestructural design criteria for the basic aircraft. The criteria presentedherein shall be utilized for preliminary sizing of the structure. Asrefinements in structural, mass and aerodynamic characteristics aredeveloped, these criteria may likewise be modified.
Results of a flutter study will dictate wing stiffness requirements toensure a flutter and divergence design without control reversal. The 15percent margin (1.15 times maximum operating speeds) required byMIL-A-8S70 (for manned aircraft) shall be considered a requirement forthe wing. The empennage already possesses this margin.
The Mach-altitude envelopes for the standard BQM-34E and the researchvehicle with the supercritical wing are presented in Figure 3-56. TheVH curve for the modified aircraft was generated by using a constantdynamic pressure curve (equivalent 'to Mach 0.95 at sea level) up toMach 0. 98, then a constant Mach number to upper altitudes. VL for themodified craft was generated in the same fashion, with the constantdynamic pressure curve corresponding to Mach 1.05 at sea level. Con-stant Mach number for V is attained at Mach 1.08.
Sea-level V-n diagrams for both symmetrical and unsymmetrical flightare presented in Figures 3-57 and 3-5S for the basic BQM-34E. Figures3-57(b) and 3-58(b) only (the two lower envelopes) shall also be applicablefor the craft wifh the supercritical wing,
Gross Weights
The gross weights for structural design of the BQM-34E are presentedin Table 3-15. The structural design gross weight for the craft with thesupercritical wing shall be the same as that for the standard BQM-34Ewith external tank (i.e., free flight ~ 2500 pounds, ground launch ~ 2900pounds, etc.).
82
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c.. Figure 3-56. Mach Number vs. Altitude
83
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SEA LEVEL
200 -100 GOO SOOEQUIVALENT AIRSPEED, Vg( KNOTS
(a). Gross Weight = 2037 Lb.
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OHK
-2200 -100 600 800
EQUIVALENT AIRSPEED, Ve> KNOTS
(b). Gross Weight = 2500 Lb.
NOTE:(b) is applicable only to the craft with the supercritical wing.
Figure 3-57. V^-ii Diagrams - Symmetrical Maneuvers, Model BQM-34E
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NOTE:(b) is applicable only to the craft with the supercritical wing.
Figure 3-58. V-n Diagrams - Unsymmetrical Maneuvers, Model BQM-34E
TABLE 3-15S T R U C T U R A L DESIGN C R I T E R I A S U M M A R Y ,
PAKACH UTE RECOVERY
CONDITION
FREE FLIGHT •
Symmetrical Maneuvers
Asymmetrical Maneuvers
Gust
CAPTIVE FLIGHT
Taxi, Takeoff, Landing
Gust
PARACHUTE RECOVERY
Drag Parachute Deployment
Main Parachute Deployment
GROUND LOADS
Ground Launch
Ground Impact ' '
Ground Handling Loads
Shipping
Hoisting
Jacking
Carting
WATER LOADS
Water Impact
DESIGNGROSS
WEIGHT(pounds)
2:;oo2037
2544
1250 to2028
1922
2900
2100
1400
1900
2944
2544
2544
1720
ULTIMATEK.ACTOK
OFSAFETY
1.25
1. 50
1.25
'
1.50
1.25
MAXIMUM LIMITLOAD FACTOR
X
±2 .50
-12.0*
7.00-
± 6 . 0
±4.0
±0.4
±0. 5
:±2. 0
±3.0
ny
±1.50
±3.0*.
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±3.0
±1.33
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±0.5
±1.33
±4.0
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1.0 to 4.0
-3. 0 to 6. 0
G. 00*
2.40'
12.0
±2.0
2.67
2 . 0 -
2.0
7.5
COMMENTS
SubsonicSupersonic
cj = 1. 5 rad. /sec. 2, basicmax
27 fps (EAS) at V
For design of attachments andsway braces. Loads act simul-taneously.
50 fps (EAS) at V of DP-2E,H
Based upon 15, 000-lb. para-chute load.
Based upon 12, 000-lb. para-chute load.
Includes JATO unit and externalfuel tank.Includes JATO unit.Ground launch loads and loadfactors based upon -JATO thrustof 14,000 pounds plus enginethrust.
n acts alone and in combinationz
with horizontal load factors.
Load factors act independently.
NOTE:. For flight with external fuel Link (2500 Ib.). V)t and Vy. are Mach 0.95 and Mach 1.05. For flight withoutexternal fuel tank (2037 Ib.), V'u and \'i. follow constant dynamic pressure lines from Mach 1.1 to Mach 2.3 and fromMach 1.2 to Mach 2.5, as shown.
* Used for equipment installation.
(1) The XIvQM-:>4F was designed for gmumi impact. .M:houp;!i the requirement for ground impact was not carriedover to the BQM-34K cr i te r ia , the s i re i iKih inhr rem w i t h the original design still exists.
Centcr-of-Gravity Envelope
The ccnter-of-gravity range for free-flight structural design conditionsshall be 15 to 50 percent of the wing mean aerodynamic chord.
Flight Loads Criteria
Free-Flight Balanced Maneuver. - Loads shall be determined at criticalpoints on and within the V-n diagram (Figure 3-57(b)) for the symmetricalbalanced condition described in Paragraph 3.2.1 of MIL-A-SSG1. Thiscondition has zero pitching acceleration, and the pitching velocity shall beobtained by solution of the expression q = (g/V^-) (113 - 1), where V-r isthe true velocity.
Free-Flight Maneuver with Specified Pitching Acceleration. - Loads shallbe determined on or within the V-n diagram for the symmetrical maneuver,with specified pitching acceleration described in Paragraph 3.2.2.1 ofMEL-A-SS61. This condition shall have a basic pitch acceleration of 1.5rad./sec.2 and the values of pitching velocity specified by Figure 2 ofMIL-A-8861. . . . .
Free-Flight Accelerated Roll. - Loads shall be determined at speeds to VL
and a.t ini t ial load factors between 1.0 and 4.0 g for the accelerated rollmaneuver described in Paragraph 3.3.1.1 of MIL -A -SS 61. The V-ndiagram for roll maneuvers is shown on Figure 3-58(b).
Control System Limitations. - The above structural design free-flightmaneuvering criteria have been selected to provide adequate marginsbeyond those maneuvers attainable in flight with an operational flightcontrol system, provided that system has characteristics similar to thaton the standard BQM-34E.
Free-Flight Gust Encounter. - Free-flight gust load factors shall be <determined by the discrete gust analysis described in Paragraph 3.5 ofMIL-A-8861. The gust velocities specified in the referenced paragraphare''unreasonably high for an unmanned recoverable vehicle. An analysiswas therefore performed on the BQM-34E flight profiles to determine amore rational value. The analysis showed that a gust of 27 feet persecond would be encountered (on the average) once in 10 sorties of themission which was determined critical (low-altitude dash mission). Thisvalue shall be used as the free-flight gust velocity at Vpj.
87
C
Captive-Flight Design Conditions. - The loads imposed on the BQM-34Eby the attachment I ' i iUn-jjs and sway braces shall be calculated using theload factors specified in Table 3-15. The gust load factors in captiveflight shall be determined using the methods and gust velocities presentedin Paragraph 3.5 of MIL-A-sti61, where the forward speeds are thoseappropriate to the launch aircraft.
Design Features Affecting Determination of Critical Conditions. - Alargc'part of structural design is governed by parachute recovery andsurf ace-impact conditions. Since the basic craft was designed for highsupersonic speeds, the lifting surfaces have small thickness-to-chordratios; this results in design of these surfaces for rigidity as well asstrength. The fact that the craf t is designed for higher speeds, dynamicpressures, and load factors than the carrier aircraft tends to makecaptive-flight conditions noncritical, except for local attachments. Aresult of these design features is to make this type of craft less criticalfor certain conditions than would be the case for a conventional, pilotedaircraft.
Elevated-Temperature Criteria. - Combined aerodynamic loading andheating were investigated for the following conditions for the basicBQM-34E:
a. Mach 0.55 at sea level.
b. Mach 1.05 at sea level.
c. Mach 2. 5 at 35, 500 feet,
These points were shown to be critical during preliminary design studieson the BQM-34E. Since the craft with the supercritical wing will notoperate outside the above points, they shall be utilized as criteria for themodified aircraft also.
Ultimate Factors of Safety. - The ultimate factor of safety shall be 1. 25for free-flight and recovery conditions and 1.5 for captive flight. •
j
Parachute Recovery Loads Criteria
Drag Parachute Deployment. - A 15,000-pound drag load shall be con-sidered for structural design during drag parachute deployment. Thisload shall act anywhere rearward within a 5-degree angle to the line offlight. A gross weight of 202S pounds was investigated for the BQM-34E.In addition, a gross weight of 1250 pounds was also investigated in orderto provide a maximum longitudinal load factor for equipment installation.
88
Main Parachute Deployment. - During main parachute deployment, largeloads are t ransmit ted to t h e - c r a f t through the forward and aft main para-chute bridles. For s t ructura l design of the 13QM-34E, a main parachuteload of 12,000 pounds shall be considered to'act in the fuselage plane ofsymmetry at angles between the vertical and directly aft. Gross weightsup to 1922 pounds apply for main parachute deployment on the I3QM-34E.
Paragraph 3.7.2.1.1 of the BQM-34E detail specification (Tcledyne RyanSpecification No. SD-2019 H-l) states that the parachute shall be suitablefor lowering the craft at a maximum sinking speed of 20 feet per secondwith no fuel abaord. Tin's is adopted as a parachute design criterion.The sink speed measured during qualification testing was 17.2 feet persecond, based on a suspended weight of 1400 pounds. This is equivalentto a sink speed of 19.1 feet per second for a suspended weight of 1720pounds.
Ultimate Factor of Safety. - The ultimate factor of safety shall be 1. 25for parachute recovery conditions.
Ground Loads Criteria
Ground loads are incurred in two separate phases: ground launch andground handling. The original criteria (for the BQM-34E) included a
^ , requirement for ground impact which has subsequently been deleted.However, structural strength for this condition, which is defined inTable 3-15, still exists.
: Ground Launch Criteria. - The ground launch weights are shown in Table3-15. The limit design thrust of the JATO unit shall be taken as 14, 000
; ' pounds.!
i Loads shall be determined for the specified conditions at center-of-gravitylocations determined by weight analysis. In addition, loads shall be deter-*mined for actual weights and ccnter-of-gravity locations for a range of
j < mission weight distributions.
; Ground Handling Criteria. - Loads shall be determined for shipping,I ; hoisting, jacking, and carting. The weights and load factors for these;; conditions arc presented in Table 3-15.
: Ultimate Factors of Safety. - The ultimate factor of safety shall be 1.5/ for ground launch (bottle ignition) and ground handling and 1.25 for
ground launch (bottle burnout).
Water Loads Criteria
Water-Impact Load Factors. - For water impact, the structural designload factors listed in Table 3-15 are 7.5 vertical, -.!:3.0 longitudinal, and±4.0 lateral. These criteria, which were used for both the XBQM-34Eand the BQM-34E, were based upon calculated values (and past experi-ence) and were demonstrated to be adequate dur ing controlled drop testsand during XBQM-34E flight operations.
Gross Weight and Ccnter-of-Gravity Locations. '- Water-impact loadsshall be 'determined for the weight shown in Table 3-15 and for the center-of-gravity location determined by weight analysis. In addition, loadsshall be determined for actual weights and center-of-gravity locationsfor a range of mission weight distributions.
Sea Conditions. - Since the craft is an unmanned vehicle descending on aparachute rather than a seaplane landing at relatively high velocities, noinvestigation of the effects of different sea states on water-impact loadshas been made. The criteria contained herein are intended to providestructural integrity for lan'ding under reasonable sea states (3 or below).
Ultimate Factor of Safety. - The ultimate factor of safety shall be 1.25for the water-impact condition.
3.3.2.2 Supplement to Preliminary Structural Design Criteria
This, supplement is to be utilized in the event that a mid-air recoverysystem (MARS) is to be incorporated on the modified BQM-34E researchvehicle. The MARS will be identical to the system incorporated on theBQM-34F; hence this criteria is developed from the BQM-34F criteria.
-- ' ,'i—
The major change required to convert the recovery system from thestandard drag-main parachutes (on the BQM-34E) to the drag-main/engagement system (on the BQM-34F) is the exchange of parachutes andattachment of the slightly longer container. Hence, the preliminarystructural design criteria for the BQM-34E re'searchvehicleareapplicableto all aspects of the vehicle operation except for MARS. The preliminarystructural design criteria for the craft with MARS are summarized inTable 3-16.
Drag Parachute Deployment
The drag-parachute deployment criteria for the research vehicle shallremain unchanged.
190;
TABLE 3-16
STRUCTURAL DESIGN C R I T K H I A SUMMARY,M A K S K E C O V K H V
CONDITION
FREE FLIGHTCompiute Target
Symmetrical .ManeuversAsymmetrical ManeuversGuat
CAPTIVE FLIGHTTaxi, Takeoff, Landing
Gust
PARACHUTE RECOVERYDrag Parachute Deployment
Main/Engagement ParachuteDeployment
HELICOPTER R E T R I E V A L
Pickup and Towing
Docking
GROUND LOADSGround Uiunch
Ground Handling I/jadsShippingHoistingJackingCarting
WATER LOADSWater Impact
DESIGNGHO.SS
WEIGliT(pounds)
25002037
2544
1250 to2023
1922
1571 to1720
29002100
19002S4-125442544
1720
ULTIMATEFACTOR
OKSAFETY
1 . 25
1.50
1.25
-
1.501
1.50- .
V
-
1.25
M A X I M U M LIMITLOAD FACTOR
nX
* •>. 50
-12.0*
00
± 1.0
7.00-
l 4.0i 0. 4
± 0 . 5i 2.0
± 3.0
ny
± 1.50
i 3.0*
0± 1.0
0
± 1.50*
i 1.33± 0 . 4i 0. 5
i 1.33
i 4. It
n z
-2.0 to 5.01. 0 to 4. 0
-3. 0 to G. 0
6.00*
2.01.01.0
2.40*
.
±2. 02.672.02.0
7.5
COMMENTS
Subsonic (with fuel jxxi)Supersonic (without fuel pod)i =1.5 rad. /sec.", basic
max
27 fps (EAS) at VH
For design of aUaeh;neni3 andsway braces. Leads act simul-taneously.50 fps (EAS) at Vf, of DP-2E.
BMScd upon 15, 000-lb. para-chute load.
Uased on test or analysis(Mininium load of 12, 000 Ib.per BQM-34E criteria. )
Maximum load factor of 2. 0acting within 45° of positiveZ axis of target.
Includes JATO un i t and fuel pod.Includes JATO unit.
Ground launch loads unil loadfactors based upon JATO t!-.r-.;stof 14, OOO pounds plus enginethrust.
•
nz acts alone and in combinationwith horizontal load factors._
Load factors act independently.
NOTE: For nigh', with fuel pod (2500 Ib.), V,, and V, are M . 61 and M 1. 05. For night without fuel pod (2037 Ib.), VH andVL follow constant dynamic pressure lines from M 1.1 to M 2. 3 and from M 1.2 to M 2. 5.
*Used for equipment installation.
91
Page Intentionally Left Blank
Main/Engagement. Paraciui l c nejrtovjmjiit. -Main/engagement parachutecriteria arc summarised in Table :i-10. Loads generated from thesecriteria shall nut be less than those for a standard (no-MARS) system.
Additional Loading Condi t ions - Helicopter Retrieval Loads Criteria
In the event that M A K S is incorporated into the BQM-34E researchvehicle, the structure shall be capable of sustaining loads developedduring helicopter pickup, towing, and docking operations.
Gross Weight sj3.nd_ Center -of-GravU.y Locations^ - Helicopter retrievalloads shall be determined for the range of.weights shown in Table 3-1G andfor the ccnter-of-gravity locations determined by weight analysis.
Helicopter Pickup and Towing. - The maximum load factor acting at thecenter of gravity of the vehicle during helicopter pickup and towing shallbe 2.0. The line of action of the pickup or towing force shall be con-sidered to lie anywhere within a cone generated at 45 degrees to thepositive Z axis of the aircraft. These criteria are based on past experi-ence and have been used successfully on other Teledyne Eyan pilotlessaircraft.
Docking. - The maximum toad factors for docking are presented in Table3-16.
Ultimate Factor of Safety. - The ultimate safety factor shall be 1.5 forretrieval conditions.
3.3.3 Structures and Weights
3.3.3.1 Wing Location Structural Evaluation
The four different structural design configurations (Paragraph 3.3.1) for <joining the wing to the fuselage were evaluated in depth. The configura-tions are quantitatively rated and ranked in an orderly fashion to helpfacilitate a design decision.
The fuselage structure evaluated reacts the wing and provides overallbending and shear continuity to the fuselage. This portion of the fuselagecontains fuel, and sealing is a consideration. The structure affectedby the trade includes fuselage skin, frames, bulkheads, longerons, andfittings in the center section region between Stations 235. 5 and 274. 59.
Other systems which may be affected by the wing location are the fuelplumbing and sealing, the inlet duct, and the electrical harnesses.
Design Alternatives (See Referenced Drawing's)
The four wing locations evaluated arc as described below:
a. Configuration 1-30-1, standard fuselage. Wing center boxwith provisions for attaching to the original wing mountingpoints on the fuselage.
b. Configuration 1-30-2, standard fuselage. Low mid wingbolted to side of fuselage. (A continuous wing would inter-fere with the inlet duct.)
c. Configuration 1-30-3, stretched fuselage. New center plugwith provisions for attaching low midwing of continuousconstruction.
d. Configuration 1-30-4, standard fuselage. Wing located ;below fuselage.
Method of Evaluation
Each configuration was evaluated7 for the following, items, called figuresof merit:
a. Weight (pounds) of airframe
b. Vehicle aerodynamic performance
c. Costs
d. Manufacturing schedule
e. Operational characteristics
Certain parameters were associated with each merit item. These para-meters are shown in Table 3-17. Merit points were assigned to eachparameter on the basis of its relative importance.
The advantages and disadvantages for each configuration were comparedand ranked in a matrix.. Table 3-18 is a tabulation of the comparison.The relative importance of each parameter and its influence on the designdecision were accounted for with this table.
931
TABLK 3-17
FIGURES OF MERIT, ASSOCIATED I 'AKAM E T K K S , WEIGHTING FACTORS
Figure of Merit Pa ram etc rs Points
Weight (100 pts)
A ero-Pe rforrnance
(100 pts)
Unit Costs (100 -p t s )
Manufacturing Schedule
(100 pts)
Operational Character-
istics (100 pts)
o W e i g h t
0 A. (Sq. F t . ) Wetted Area
o F . R . Equiv. - body f ineness ratio
o V f.j Horizontal Tail Vol.
c V y Vert ical Tail Vol.
•<b Fabricat ion Complexity*
o Quality Assurance*
o Producibility*
0 Abi l i ty to Hold Tolerance1
0 Fabrication Complexity
o Geometric Restr ict ions
0 Quality A ssurance
o- Abi l i ty to Hold Tolerance
o Producibil i ty
$ Susceptibil i ty to Damage
0 Maintainability, Repairability,and ease of field assembly -
o Reliability
O Safety
100
25
25
25
25
Total 100
35
25
5
10
35
40
10
20
t ^Parameters which a f f e c t more than one f igure of merit
94
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The final evaluation of each configuration was based on the number ofmerit points earn eel. Kach point is a mark against the configurat ion.The candidate design with the lowest total was considered the best selec-tion. Table 3-19 is a summation of points for. each configuration.
x
Final Selection
The evaluation of the alternative configurations is summarized as follows:
a. The high wing, configuration 1-30-1, scores best, except forthe fact that.this location is not typical for transports.
b. Fabrication of a new center plug and stretching of the fuse-lage are quite expensive relative to the other configurations.The increased welted area gives it a poor aerodynamic ranking.Fabrication complexity and impact on other systems rank highrelative to the alternative configurations.
c. In configuration 1-30-4, the bottom of the fuselage does notscore well acrodynamically and ranks third from an opera-tional standpoint.
d. Configuration 1-30-2 is the most feasible wing location,based on the total evaluation of all items.
3.3.3.2 Structural Analysis
A structural analysis was performed on the fuselage center section tosubstantiate the feasibility of mounting the NASA research wing to theBQM-34E fuselage. The internal load distribution generated by the wingreactions on the fuselage was determined, and an estimate was made ofmodifications required.
For analysis purposes, the fuselage center section between XF 23.3.5 and •XF 274.59, where the wing is located, was isolated in a structural model.The. wing introduces large concentrated loads, at the attachment points,which are required to be distributed into the shell; In addition, this sec-tion provides overall bending and shear continuity to the fuselage. Italso affords fuel containment and is subjected to fuel pressure.
Configuration 1-30-2 (wing bolted to fuselage side) was analyzed in detail.The structural concept requires a two-piece wing, consisting of a leftand right-hand panel. Wing continuity or carrythrpugh structure is
98
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provided by the fuselage, which is modified to accommodate the wing.This modif ica t ion consists of the following items:
a. Three new frames, which serve as the main carrythroughstructure.
b. A left and right-hand wing attachment fitting. These fittingscomprise the structural link between the wing and thefuselage.
c. The tic bars which connect the ends of the horseshoe-shaped frames and provide'fuel-containing facilities toreplace the original wing structure which performed thisfunction.
Construction consists of a conventional semimonoc.oque frame/longeronshell structure. The material is aluminum alloy. Drawing No. 1G6SCW014depicts the configuration.
The fuselage structure in the region of the wing attachment is idealizedinto an analysis model which describes the geometry, reactions, materials,and the structural elements. The finite elements include rods, bars, andshear panels. These elements and the computer program formulationsare described in Reference 11. The program computes displacements,reactions, and internal loads on the finite elements. The wing reactionsare applied as concentrated loads at the attachment points. The portionof structure between XF 233.5 and XF 274.59 is isolated into the analysismodel (Reference Drawing No. 1GGSCW014). Appropriate reactions areprovided at these stations.
Input
The structural analysis input data consists of the following items:
a. Geometry
b. Idealization '
c. Structural elements and sizes
d. Material properties
e. Reactions (constraints)
f. External loads
These items are described and presented on the following pages for thedifferent structural configurations used in this feasibility study.
100
Geometry. - Center Scctilion
XF 233.5
1GGF002 FuselageCenter Section
n
Inlet Duct '
A
XF 274.59i
-_^,S^.-^ _ -Jf_ - J~> 14-1 7- - 7n ffi Hi N
Wiiut Root
New Frames(3 Places)'
Frame Tie Bar(3 Places)
Wing Fitting
Fuselage Fittings
SECTION A-A
101
Grid Points and Shear Panels.
10 TOP
19
X0,0
Typ. Coord.X, Z
102
Rods and B:irs.
53
103
Rods and Bars.
19 36
Sta. 15.5
30
44
937
43
104
Analysis - Configuration I-:iO-2
External Loads. - Loads representing the 5-g symmetrical maneuverflight condition are applied to the center section. Wing bending and shearloads are distributed equally to the three fuselage frames. Torsion isreacted by the frames at the forward and aft spars.
Fuselage bending and shear resulting from loads.applied on the forwardfuselage are applied at XT 233.5. The section is fixed at XF 274.59.
X 233.5 X 274.59r
Wing Root Loads (Ultimate)
Fuselage
M = 379,332 in. Ib. (u l t . )T = -15,600 in; Ib. (ult .)V = 8,130 Ib. (u l t . )
Resolve Loads Across Fuselage
Condition; Sym. Balanced5g Maneuver
105
M' = M cos 6 ~ T sin 6= 370:332 cos 38.95° - 15600 sin 38.95°= 28519G in . lb . (ul l . )
T' = T cos 0 + M sin 9
= 15GOO (0.77769) + 379332 (O.G2SG4)
= 250595 in.lb. (ult.)
V} - S130 Ib. (ult.) /
jLoads - Win;.;- Reactions onFuselage^ - Assume M1 and V1 are sharedequally at throe i'rames.
-, /T, ;U _i.-;.u.yuM/I-rame = — =
<J (J
= 95005 in . lb . (ult.).
T, /r, V 8130P /Frame = — =V 3 3
= 2710 ib. (ult.)
M/frame is coupled into the frame as shown below.
V
95065M = 36563 Ib. (ult.)
Torsion, T', is coupled between the forward and aft frames.
106
Torsion
Fuselage
/ H H -|1 1 1
PT
1 \PT
T'12.4
250595T 12.4 12.4
= ±20209 lb. (ult .)
SUMMARY OF APPLIED .WING LOADS ON FUSELAGE (ULTIMATE)
GRID "POINT*
21 (25)22 (24)
30 (34)
31 (33)
39 (43)
40 (42)
X(lb.)
36563 •
-36563 .
36563
-36563
36563
-36563
Y
000000
Z (lb.)
_
22920-
2710
-
17500
*See Idealization(25) indicates opposite hand grid point
Forward Fusejagc Loadsf
!XF 233.5 shear and bending (limit)
ii M = -170645 in.lb.
V = -3015 lb.
5-g sym. maneuver
107
The loads are increased to ult imate and panel pointed as shown below,
30ir> li).
8120 lb.
16,252 lb.
Reactions
The section is constrained at XF 274.59 at grid points shown.
R
Material Properties•X 274.59
Material is 7075-T6 al. aly.
= 73000 psi
F =cy
F =su
E =
G =
65000 psi
43000 psi
10.5 x 106 psi
3.9 x 106 psi
108,
Output. - The results of the computer analysis include the following data:
a. Internal forces and stresses in bars, rods, and shearpanels.
X
b. Deflections
c. Reactions
These results are summarized and presented in the following pages.Only the significant loads are shown. The detailed output for each elementis included in Reference. 12.
Output - Longeron Loads. - Critical longeron loads are tabulated below.
16GF295 Roof
16GF115Shear Web
1GGF269 Fitting
166F268
166F260 Keel
LONGERONPART NO.
166F295
16GF268
166F2GO
ELEMENT NO.
334
339
361
LOAD(Lb., ult.)
19317 Ten.
15266 Ten.
-1G252 Comp.
109
Skin Shear Loads. - Crit ical skin shear loads are shown below.
ELEMENTNO.
1017101810191025102G10271028
STRESS(psi, al t . )
1192G2G65917150112373G7G220728
2416
. SHEAR FLOW(Ib./in. , ul t . )
5961333856562
18381036121
Frame Loads. - The forward wing attachment frame internal loads areshown below. Loads are symmetrical about the vertical and centerline.
ROD NO.
390517518519520
M (ult.)(in./lb.)
-36741-565904856812000
axial'(lb. , ult.)
-4S75
9628-531411189
SHEAR -(lb., ult.)
46003160
-291255299
110
Reactions
U l t i m a t e reactions are shown below. Loads are symmetrical about the
vertical center l ine.K3
GRID POINTNO.
1. 9
o
45
R 2 (u l t . )(pounds)
-13715-1G150
GOSG161 0015348
R3 (ult.)(pounds)
-1671-2374
-923-183 •
73
Output - Deflections. - Displacements along the bottom and based onultimate loads, are shown below.
GRID POINTNO.
5142332415050
DISPLACEMENT(inches)
0.0-0.070-0.202-0.424-0.682-1.010-1.698
111
Stress Analysis
A stress analysis was made on the c r i t i c a l s t ruc tura l elements to deter-mine t h e i r ability to carry the loads shown on the preceding pages. Themost crit ical elements :uv certain panels that comprise the fuselage shelland the frames which react the wing loads. The skins undergo severediagonal tension, and it is recommended the 0.050-inch basic skin bereinforced with an O.Oi:Mnch doubler over several panels. The framesexperience large bending loads and must be adequately stiff to prevent
large wing deflections.
Part No. 1GGF2GS Longeron
Rivets
, ZO.Typ.
.390
Tension across net section:
P'- 15266 Ib. (ult.)
A = 0.542 in.2
f. =152GG
t A 0.542
f = 28166 psi
F = 76000 psitu
_MS =
7600028166
1,12
- JL. JLJ.
2
0.20
2
ro
tw
2
O o n. £U
The tension fit t ing end of tin; longeron is analyzed as an angle-type liltingwith a NAS62G bol t , usiiv rn-lhoH.s of Reference 3.i { • * • * • • - •
A = 1.55 -^ = 1.45 y. = ~ = 0.18752 l 2
B =.- i oi[ _^i±^ = 1.11 y = •6 4 G - 0.322JJ • * • • — - * • rt ^v O
C - 0.61 -
—~ =•• O.S15 Ag = * a tw = 0.512
d = a- F-- - 0.3005 c = 0.637a - 0.5191
I = 0.298 a3 t = 0.0322\v
Wall Stress
Axial load, f :tu
f = — = --—- = 31543 psi (ult.)tu Ag . O.ol2
Bending stress, f :
M •= P (c-b) = 1G150 (1.5191 - 0.3005)
M = .3530 in./lb. (alt.)
f = c = - 0.407bu I 2
3530 (0.407).bu ~ 0.0322
f = 44618 psi (ult.)bu
113
Net Stress:
81 315'13 + 44613
- 76IG1
bu
i (U{L)
1.25 = 95000 psi
MS «
Pad Stress, f .buo'
= 0.230 a -c!~y~~ - 1.599
/ •
o
= 0.55
fbuet
bue
k"
psi
= 76000(1.5) ='
0.39
•) 1". r-1.95
MS =91103 lp.25
Wing Attachment Frame. - The frame experiences critical loads at the
wing attachment.
1.4 ~1
• T•) 9 1-5°
1.0
Loads, (ult.) @ A-A:
M = 56590 in./lb.
P = 9G2S Ib. (axial)
V = 3160 Ib. (shear)
Bending @ A-A:
Sec. A-A
I = 1.23 in.4.
A = 3.1 in.2
.1 A - '" 1.23
f = 53715 psi (ult.)
F = 76000 psitu
56590 (1.1) 9C28
3.1
115
_ _ _ - c l i - ~ T n c kecl o r i iv ina l ly -served t o distribute t h evertical reaction of. the external fue l t ank to f rames and to reaet bendingloads. Its fuhelion in the NASA research vehicle is primari ly as alongeron for fuselage bending loads.
Geometry:
r -M-- - i ti .
1II
-»— 1 . 74 1-
f
\
JI— .75-—
i.030
Material: 7075-TG51Area = . 64S in. 2
.050
Material = 7075-T651
Area = 0. 648 in.
Maximum compression, fc:
P =-16252 Ib. (ult.)
= — = 1G252 ' .''c A ~ 0.643
f = 25080 psi (ult.)C
Allowable compressive load, Fccr
Kc
E I \2/ 1 \ccr 12 (1-u1-)
K = 4.0 for center
K = 0.43 for !'huiL',e\j
Reference 4; Table 7
b -- 1.7-1 in. for center
b - 0.75 in. for flange
u ---• O . :>0
E = 10.5 x 106 psi
t = O.OS in.
"> G /4 ft"j0.5 x 10 / O . O S
11774cr 2center 12 (1 - 0.3 )
= S0247 psi > Fcy
cyflange
Margin of safety:
0.43 772 10.. 5 x 1QG
12 (1 - 0.32)0.03
1 0.7-5= 46430 psi
4643°250SO
0.851
Skin Panels. - The skin panels which comprise the fuselage shell experi-ence high shear flows during the 5-g symmetrical maneuver. Theseloads are mainly due to coupling out of the wing torsion into the shell.
Geometry:
Material 7075TG Clad
5.0
R = 12.5 it = O . O G 2 i n . (cloubler added)
117
Loads; elements 1026 and 1018 are critical;
- q 1020 ~ l8;-!8 Ib./in. (ult.)
q 1018 = 1333 Ib./in. (ult.)
Shear stress, f :
~ .s t " 0,062
f = 29645 psis
Shear buckling allowable, F
ft\2.•F = K Escr s \b
52 = • , _. r~ 3 'ao
Rt 12.5 (0.062)
.',K = 14 (Reference 5, page 396)S
E = 10. 5 x 106
Fscr
F = 22638 psiscr F
./.semitension field
A. - * i 31F 22638
scr
,'.k - 0.12 (Reference 5; page 407)
118
x:
_ta = 0 .062(6) =
Ae 0.542
tana = 0.98 (Reference 5; page 408)a = 44°25'
F = 32800 psi (panel allow, Reference 5; page 410)s\v • /
Panel margin of safety:
29645.
3.3.3.3 Structural Configuration
The structural configuration selected for the NASA research wing isshown-in Figure 3-54. • The wing consists of the following items:
a. Wing structural box
b. A machined root rib
(' c. Leading edge'
d. Trailing edge
6; Wingtip
f. Leading and ti'ailing edge flaps
g. Ailerons
The wing "box is of full-depth sandwich honeycomb construction. Itstapered skins are adhesive bonded to the core. Lightweight sheet metalspars bonded to, the skins and core form the spanwise sides of the. box.The root rib is attached to the inboard ends of the skins, core, and'spars.Provisions are at the outboard ends for attaching a fiber-glass wingtip.The leading and trailing edges attach at the spar flanges, flush with thecenter-box skins to form a smooth, aerodynamic surface. Movable sur-r
. faces are appropriately hinged from the leading edge spar and frameshoused in the trailing edge. The wingtip and trailing edge are of fiber-glass construction. The remaining wing structure, including the honey-comb core, is of aluminum alloy consti'uction.
119
c
The movable surfaces are actuated by hydraulic units located in theleading and t ra i l ing edges. Appropriate pushrods, bell cranks, andfittings link each movable surface to its actuator. A wing fillet fairingis provided. /
A qualitative evaluation of the configuration indicates that the weight-cost comparison relative to other concepts is good. Fabrication com-plexity is not great, and susceptibility to damage compares with alterna-tive designs. Teledyne Ryan Engineering and Manufacturing have hadmuch experience with this type of structural configuration as fabricatedfrom both metallic and advanced composite materials. Technical risksare not high.
The configuration lias a multiplicity of load paths, and any local damageis not likely to affect surrounding structure. Wing bending, shear, andtorsion loads arc carried by the center box. The root rib redistributesthese loads and reacts them into the fuselage attachment bolts. Loadson movable surfaces and leading and trailing edges are distributeddirectly into the wing box. Loads on the actuating systems are notsevere. Adequate space exists for installation of the actuator systems,and ease of accessibility is provided.
The concept offers an aerodynamic surface with a high degree of smooth-ness and lends itself to fabrication.
3.3.3.4 Mass Properties
The weight and balance for the NASA wing feasibility study (low wing) ispresented irv Table 3-20. The base vehicle is the BQM-34E (Reference 7),modified to include the new wing, control surfaces and controls, MARSsystem and repositioning of the wing and ai-ea rule fairings. The pay loadconsists of a shaker installation and available volume located at Body.Station 213. The density used for this volume is 45 pounds per cubicfoot. Additional allowances for payloads are covered in Paragraph 3.3.2.
The center-of-gravity travel is considered to be for level .flight and to belinear from a zero-fuel-weight configuration (29. 93 percent MAC) to agross-weight configuration (10.56 percent MAC). With the presentsystems and payloads, no ballast is required. If any equipment is modi-fied, replaced, or removed, further study should be made to determinethe effects on the center-of-gravity travel and the possibility of ballast.
120
c
TABLE 3-20
WEIGHT AND BALANCE SUMMARY,NASA WING STUDY (LOW WJNG)
Aorocyn-'inic SurfncesWing (Key incl.
Fairing)Fin (1)Stabilizer (1)
SacJy"KOSC (1)Center (I + Mod)Tail (1)
Take-Off & Recover"Take-Off (I)Recovery (2)
PropulsionAir Breathing Sys(lFuel & Lub (1+MoJ)
Power Generating SystemElectrical -- AC(i)Electrical - DC(1)
Orientation ContrStabilator & RudderLeading Edge Flaps'
(New)Aileron (New)
Guidance & Electronics(New)
Environmental Protectio;(I)
Hydraulic System (New)
Payload (Nesv)Shaker Inst'lAvail @ Body Sta 2i:
Area Rule Fairing (New)ForwardMid
V.T.
(1)
(+5.0)+ 5.8
00
(+.12.7)0
+12.70
(+43.7)0
+43.7
(+13.1) '0
+13.1
( 0 )00
(+42.0)0
+ 14.0
.+28.0
(-41.0)
( 0 )
(+20.0)
(+119.9)+ 95.1+ 24.8
(+ 23.3) .
WEIGHT
(Lbs . )
(214.10)162. ',0
31.7319.97
(271.61)101.02
• 115.3955.20
(i'65.23)6 . So
153.40
(498.20)443.4054 '.80
(125.40)12.20 -.113.20
( 77.80)35.8014.00
28.00
( 39.00)
( 19.70)
( 20.00)
(119.93)95.1324.80
( 23.3)14.45.0
HORIZONTAL
ARM
(In.)
(290.6)270.9
3-'.6.-i362 . 3
(2-VS.3)181.91&. .2329.4
(37J.5)257.4
• 333.3
(237.1)291.5251.7
(219.8)184.0223.6
(316.0)353.2276.8
233.1
(171.7)
(285.9)
(240.8)
(166.1)153.9213.0
(260.6)205.0312.0
MOMENT
(In.-Lb.)
(52223)43994
109907239
(670A3)1837830434 .13131
(62563)177160792
(143042)•129250 '13792
(27.560)224525315
(24587)126453375'
8067
(15243)
( 5632)
( 4816)
(19923) "146415282
( 6072)29521560
VERTICAL
ARM
(In.)
(47.2)"1.0
73'.157.0
(51.6). 54.2
40.054.1
(56.9)60 . 3 .56.7 ..
(49.3)49.443. 4
(55.2)54.655.2
(48.5)57.341.0
41.0
(53.5)
(50.8)
(49.8)"
(53.3)54.053.0
(55~.0)55.055.0
MOMENT
(In.-Lb.)
(10116)6653
23201135
(14005)547155432986
( 9405)4.158939
(24546)21o962650
( 6917)6666251
( 3772)2050574 '
1148
( 4761)
(1000)
( 996)
( 6451)5.1371314
( 1281) .792275
121
TABLE 3-20 (Continued)
WEIGHT AND BALANCE SUMMARY,NASA WING STUDY (LOW WING)
.
A f t
Miscellaneous (1)
Weight Empty
Unusable FuelMain (1)AuxM (2)Aux//2 (New)
''nusable Oil (1)Oil - Usable (1)
Zero Fuel Weight
Fuel (JP-5) Gal .Main (1) 38.7
. Auxtfl (2) .10.3Aux//2 (New) 6.0
Refr igerant (1)
Gross Weight
LEMAC =261.57MAC = 26.70.
WT.
(1)
0
(+244.5)
(1.1)0
+ .7+ .4
00
(+245.6)
(+110.8)0
+ 70.0+ 40.8
0
(+35G.4)
reiCIIT
. (Lbs.)
3.9
( 3.30)
(1627.02)
( 5.1)4.0
o 0.70.4
( 9.3)
(1644.32)
(373.8)263.0
70.040.8
( 8.3) .
(2026 .42)
HORIZONTAL
ARM
(In.)
400.0
(169.4)
(269.89)(31. 1G%)(<:
(241.6)272 .0199.8260.2
(299.0)(221.5)
• (269.56)( 29.93%)
(244.7)254.2199.8-260.2
(128.8)
(264.39)
(10.50%)
MOMENT
(In.-Lb.)
1560
( 559)
(439268)
( 123.2)1038
140104
(688)( 2060)
(443248)
( 91457)663551393610616
( 1069)
(535774)
VERTICAL
ARM
(In.)
55.0
(112.7)
(51.38)
(42 .9)40.550.055.4
(43.0)(43.0)
(51.29)
(50.1)48.752.656.055.3)
(51.09)
MOMENT
(In.-Lb.)
214
( 372)
(83622)
( 219) •. 162
3522
. ( 90) -- ( 400)
(84340)
(18730)12303
36822240
( 459)
(103529)
- *
1 .(1) Report No. TRA 1GG44-22, Actual Weight Report for BQM-34E Supersonic Aerial Target,
Serial' No. BQ-161S1, dated 4 May 1972.
(2) Report No. TRA 1C644-25, Actual \Veight Report for GQM-34F Supersonic Aerial Target, dated4 February 1972.
(3) Percent mean aerodynamic chord (MAC).
c
12?
3.3.3.5 Advanced Composite Components
The capabilities of advanced composite materials in air-frame structurescan be demonstrated by the application ol these materials to small com-ponents, such as a skin panel, flap, aileron, rudder, or part of anempennage. The development of advanced composite technology hasprogressed along- these l ines, and many do's, don'ts, and warnings haveevolved from these types of application programs. The technicalapproach to achievement of a design objective with advanced compositematerials may he summari/.ed as follows:
a. Design/analysis
b. Fabrication
c. Test
Design Analysis
This phase of the technical approach begins with a structural configurationand a design criterion. From this, geometry, external loads, sizes, andadvanced composite material properties are generated. Figure 3-59shows an outer wing panel fabricated from composite'materials byTeledyne Ryan. This panel is a component on Teledyne Ryan's AQM-34Rdrone. An automated, iterative, design/analysis procedure shown inFigure 3-60 was used to substantiate the-outer wing panel design. Theprocedure involved three separate but coupled types of analysis: laminate,structural, and flutter.
Each analysis required a computer program. The analysis cycle involvedan inner strength loop and an outer flutter loop. The inner loop startedwith the laminate analysis program, which was used to generate stiffnessmatrices for the plate elements used in the structural analysis program.. .Ply orthotropic material properties, ply orientations, and allowablestrain data were part of the input. Information from, this phase wasused as input in the structural analysis program to determine internalloads. The internal loads on the finite elements were then cycled backinto the laminate program and were used to perform a point stress analy-sis of the laminate. Each ply in the laminate was analyzed for its criticalfailure mode, and margins of safety were calculated.
The structure was idealized into an analysis model which described thewing geometry, reactions, materials, and the structural elements. Thefinite elements included rods, shear panels, and plates. The composite
123
PUD 49-3/CE 3305Outer Wing Panel
Panel Area20.5'Sq. Ft.
WRP
Mil | I |U— 3.3 Ft.—-1
0.3 Ft.
c Figure 3-59. Model AQM-34R Drone
124
Ply Properties , Ply Orientations
StiffnessRequirements
(FLUTTERANALYSISTEA 3388
LAMINATEANALYSISTRA 1517
(SQ5)
IStre
Requir
\LiPr
ngthements
j-minateoperties
| External Loads
<i /STRUCTURAL
ANALYSISTRA 1500/XT » «"«T»T"K l -VT\(NrtOinrtrs)
Stiffness Mass- TRA 1012
Final Design
Figure 3-HO. Analys i s Approach and Cycle
125
laminated skins were represented as orthotropic triangular membraneelements. The program computed displacements, rcacliony, and in te rna lloads for the design load condi t ions. The loads on Ihc composite skinswere cycled i n J o Die l amina t e analysis program, and a point stress analy-sis was performed. Figure 3-01 shoxvs the structural analysis procedure.An elastic:-axis beam representation of the structure was used in theflutter analysis loop utili/.cci. Orthotropic beam elements were employedto de te rmine El and GJ stiffness properties.
Detail analyses were also performed on adhesive bond lines, joints, andstability failure modes.
Fabrication
This phase of the technical approach includes the following broad items:
a. Materials and test program
(1) Characterization of a resin system
(2) Tests to determine the mechanical properties of theiiiaterial system to establish design allowables
(3) Development of adhesive data
(4) Development of tests and specifications
b. Manufacturing and quality control
(1) Fabrication techniques-• • s- • •
(2) Autoclave versus vacuum bagging
c. Tooling development
(1) Structural concepts - tradeoffs
(2) Tooling materials
d. .Manufactur ing engineering
(1) Tolerance requirements
(2) Bonding processes
126
I N P U T
1. S t r u < . - U : r a i Mca l i / . a i ion2. \Vi!>i; (J ey i;; !.• try3. M a l a r i a l Proper t ies•1. h 'K-s i - iL - i i t Si/A-5\>. Constn'i i i i l - j
ComputerProgram 1G!7
E L K M K N T I M C(Plate SiiU'
ComputerProgram 1500
S T R U C T U R A L A N A I . V S L 3(Triangular Plate li
1. In te rna l Loads2. Displacements3. Reactions
ComputerProgram 1517
LAMINATE ANALYSIS
1. ' Lamina Stress2. Fa i lure Mode3. Mari r in of Sai'etv
DETAIL STRESS ANALYSIS
1. Face Wr ink l ing2. Intcrccll Buck l ing3. Core Shear4. Joints (Program LAP*)
EXTERNAL LOADS
Figure 3-G1. S t ruc tu ra l Ana lys i s Procedure
127
(3) Dr i l l ing , cu t t ing , trimming, and routing
(')) Assembly techniques
c. Quality Control
(1) Rece iv ing and inspection meetings
(2) Nondestruct ive test methods
(3) Process control
(4) Fabrication control
Tests
Component testing involves static and flight test programs.
The static test is conducted to determine the flight-worthiness of thecomponent. The test substantiates the strength and stiffness integrityof the component. The test may or may not be carried out to failure.
A flight test program is conducted to evaluate the environmental effectsunder actual flight conditions. The need for protective environmentalcoatings can be determined, and any evidence of excessive deflectionsor structural deterioration.can be noted.
Representative panels of a wing or fuselage shell, such as stringer,plates, and skin-stringer combinations, are subjected to compressiveand shear loads which demand consideration of their behavior in thedesign loading ranges. Those structural elements may be tested withconventional laboratory equipment with the use of conventional testingtechniques. Initial buckling data, overall panel stiffness, ultimatestrengths, and failure modes may be obtained and correlated withtheoretical predictions.
Teledyne Ryan has had considerable experience in the development andapplication of advanced composite components for supersonic drone air-craft. A boron horizontal stabilator for the BQM-34E (supersonicFirebce II) was designed, fabricated, and ground tested. Later, TeledyneRyan designed, fabricated, and tested three ultrahigh-modulus, graphite/epoxy, horizontal stabilators. One unit successfully passed static anddynamic tests. The remaining surfaces were flight tested at Point Mugu,California.
128!
The application of advanced composites to the BQM-34E horizontalstabilators' rc.suItecl in a -10 percent weight reduction for the boron u n i tand 50 percent weight reduct ion fur the graphite component over theexisting metallic component. This, in turn , permitted a reduction inballast weight located in the vehicle nose. The thin aerodynamic surfaceson the BQM-34E arc s t i f fness-cr i t ical ; hence, the advanced compositematerials could be used e f f i c i en t ly .
One advantage of flight test ing components fabricated from new and untriedmaterials pa an unmanned vehicle, such as the BQM-34E, is that theconsideration of pilot safety is not involved.
3.3.3.G Variable Stiffness
Concepts for varying the wing bending and torsional stiffness may beclassified in three board categories:
a. Mechanical {arrangement of structural elements)
b. Material changes
c. Combination of mechanical/material .
Bending stiffness, El, and torsional stiffness, GJ, are manifestationsof a mechanical/material system integrated into a structure. E and Gmoduli arc associated with the material, thickness, etc., implicit in thestructural arrangement. Quantitatively, I and J are expressed as follows:
Ai = Area of bending cap material
Yi = Distance from neutral axis to centre id of Ai
A = Enclosed area of a torque cell
/els^— - Line integral around the periphery of the cell walls ofthickness t
129
Concepts for mechanically varying win;; s t i f fness of necessity involvethese parameters.
Figures :<-f>2 through .'5-(5S i l l u s t r a t e concepts that will achieve variationsin bending and torsional s t i f f n e s s . The fol lowing constraints are assumedto be common for all the rncihixls:
a. The wing plani'onrv and aerodynamic shape must bemaintained.
b. The leading and trai l ing edges, flaps, ailerons, and con-trol systems remain unchanged.
c. The external loads are the same (same strength require-ments for all concepts).
d. Stiffness variat ions are achieved with the wing box.
Mechanical Methods
The following schemes are included in this approach:
a. Variable spar cap and/or stringer area.s (removable slugs)
b. Skin covers , replaceable, with different thicknesses
c. Variable torque box size
d. For wings with many shear webs or stringers, a mechanicalmeans for deactivating these elements to become structurallyineffective
Eemovable spar cap slugs influence the bending stiffness two ways. Asthe area changes, the distance between its centre id and the bending axischanges. Increased area causes an increase in centroid distance, witha. cumulative effect on the area moment of inertia. Wing mass willchange; however, its distribution can be controlled by selective areachanges. Inertia effects on aeroelastic characteristics should not differwidely.
Skin covers which can be replaced by others of different thicknesses willinfluence torsion chiefly. The line integral ~J will vary with a change int. An increase in t will give, a corresponding increase in torsion stiffness.Inertia d is t r ibu t ion will not change significantly.
See Del ail
Interchangablc InsertsFull Length of Spar,"Variable Size and/or Materials
Spar Caf .
lute i'ch an gable PanelsBetween Spars, VariableThickness and/or Materials
Figure 3-62.. Concept 1, Variable Wing Stiffness
Shear Webs
»Par Win«-£?
Conceptp-
'•riO
O
3
CLoCJ
OO
to
CO
0>
Si
133
boc
OO
OO
LT5
ICO
OS-.
CO
134;
VI
oc
o
co
oO
iCO
o
135
• Sec Detail
Removable Cover Strip
..Non-Load CarryingSkin is Fitted When InnerPanels are Used.
Variable Thickness-I Cap Strips
Inner Panels Used toReduce Torque Box Area
Figure 3-07. Concept 6, Variable Wing Stiffness
Spamvise Stringers,See Joint Detail
Ski aJoint is Detached byRemoving Bolt
DETACHABLE STRINGER JOINT
Fi-ure 11-tiS. Concept 7, Variable Win- Stiffness
IT?
Changing of the enclosed'area of the torque cell is a very effective way tochange the torsional . s t i f fness , which is a function of area squared. Thisscheme, in conjunct ion wi th changing of the cell wall (skin) thickness,will yield a wide var ia t ion in stiffness. Dcactivation. of structural ele-ments is another approach tantamount to removal of .stringer areas andchanges in torque cell si/.c.
The mechanical methods vary in cost, weight, fabrication complexity,ease of assembly, and reliability. This is also true for development andtest programs required to substantiate the concepts.
Material Changes • .
The use of advanced composite materials to achieve a variation in struc-tural response cannot be overemphasized. The orthotropic properties of.these materials make them superior to isotropic materials (metallics) totailor a structure to specific strength and stiffness requirements. Theserequirements can be controlled through selection of the materials andlamination patterns.
Advanced composite materials offer four sources of design freedom whichmay be utilized to tailor any desired stiffness. These sources are asfollows:
. a. Material selection
(1) E-glass-epoxy
(2) Graph itc-epoxy
(3) Boron -epoxy
b. Lamina (ply) orientation
c. Lamina thickness
d. Number of plies (lamina)
Figure 3-69 illustrates the wide variation in El and GJ properties thatmay be achieved in a design. These curves are for a wing componentdesigned by Tcledyne Ryan. All of the curves will meet a commonstrength requirement. Figure 3-G9 indicates that El can be varied from12.5 to 92.5 \ 10G at the wing root. The weight change is not significant.
138
Figure 3-G9. Wing; Stiffness Variation, Model 147 TF Composite Skins
139
Structural conf igura t ions inlo which advanced composite materials mayreadily be integrated \vi lh cont ro l led .sli'ffncss properties arc as follows:
a. Wing box f u l l -depth honeycomb core construction withadhesive bonded composite skins.
b, Removable skin covers wi th honeycomb panels fabricatedwith composite facings.
The tailorability and versa t i l i ty of composite materials permit the design ofa. structure for specific s t i ffness characteristics and strength requirementsunmatched by other mate r ia l s .
3.3.4 Stability and Control
3.3.4.1 I^iT^u^uwl_Ch;ir;i<.: i eristics
A preliminary analysis of the research configuration (1-30-2) equippedwith the lugh-aspect-ratio, supercrit ical \ving has been conducted todetermine the estimated static longi tudinal stability level and trim capa- •bility. The results indicate that this configuration, with the existinghorizontal tail, should have approximately the same stability margin asthe NASA full-scale f l i gh t research configuration as well as adequatecontrol power to trim the lift coefficient up to approximately O.SO.
Data and Method
The data available for the study is unpublished. It consisted of a plot of
CmCL *or *'1C wing~DO<-'y ant' wing-body-tail (Figure 3-70). Tiie wing forthe research configuration was assumed to be an exact scale of that of aNASA wind tunnel model, so that the aerodynamic coefficients for thewing could be applied directly. No corrections were applied to accountfor differences in fuselage characteristics, and rigid aerodynamic datawere used throughout. The static stability of the research configurationwas estimated on the basis of this data. - The trim require-ments at the higher lift coefficients were evaluated by the use of the windtunnel data, since the pitching moment data arc nonlinear with increasinglift ; coefficient .
iiLongitudinal Stabi l i ty
Wing-Body^ - The l i f t -curve slope of the hori/.ontal tail of the NASA modelwas calculated f rom the tail incremental stability contr ibut ion
and subtracted from the measured l i f t -curve slopes of
the wind tunnel test data to determine the wing-body' l ift -curve slope.The wing-body pitching moment derivat ive Cni -....„ was then determinedfrom <
Horixonfol Tail . - The horizontal- tai l contr ibut ion to CniQ, was obtainedby correcting; standard 13QM-;ME data 'in the presence of the body for thegeometric changes of the new wing.
Down wash slope, de/da, was calculated by available empirical methodsto be 0.31 for low speed. Calculation of-dowmvash by the same methodfor the NASA model tail location was approximately o percent larger thanfor the BQM-.'ME tai l location. The low-speed dowmvash v/as modified bythe l i f t -curve slope ratio to obtain the dowmvash as a function of Machnumber. For the NASA configuration, de/dor at Mach 0.90 v/as calculatedto be 0.46, An indop-oiicient check based on stabi.li/.er incidence effective-ness yielded 0 .4G. The accuracy of this correlation is probably for tui tousbut is nonetheless encouraging. A value of horizontal tail dynamic pros-sure ratio of 0.90 v/as assumed and applied with (1 - dc/da) to obtain thehorizontal-tail stability contribution in the presence of the wing.
Neutral Point, -r The static, stability margin was then calculated from
Cmtt WBT
\^s ~~ -• - --" ~ ;
mC °LL a WBT
and the neutral point from
No' = 0.25 - Cmr.
' L
The calculated stability margin is shown, in Figure 3-71 in comparisonwith that for the NASA model. For similar tail volume coefficients (0.925for the BQM-34E versus 0.91 for the NASA model) and for similar verti-cal displacements of the horizontal tail relative to the \\ing chord plane,one would expect similar stability levels.
The allowable center-of-gravity range is also shown in Figure 3-71. Forconservatism, the most aft center of gravity was established O.Oac" for-ward of the most forward neutral point.
The most forward center of gravity was established for the condition of-10 degrees elevator tM'ieclion and a t r i m l i f t coefficient of O.SO. Use
-••.6?. '• l.'u !
•
Vr
f'/.•-:- •
1.]
i .-
sr-'Tt.?,:
..i_L_L._; \........; ; i •: I
. I 1 j. . : ..
- P-9 r'.'/v. A'v ~A
^
Figure 3-71. Kst inuUoci SUUic S t a b i l i t y ant! CG Hai i^
was made of the wind tunnel data plots in these calculations because ofthe nonlinearities in the pitch inf.1; moment da la at high l i f t coeff ic ients .The allowable cenler-of-gravity range shown in Figure .'3-71 is shiftedapproximately 10 inches aft of the standard BQM-.'MK range, but ballastrequirements should be alleviated by the more rearward wing location.
jj;im •Characteristics
Elevator deflections required to tr im arc shown for a nominal center-of-gravity locution of 0.25c in Figure 3-72 as a function of Mach number andlift coefficient. The nonlinear pitching moment data were again used inthese calculations. No corrections were applied to the pitching momentdata due to configuration differences because of the similarity in stabilityand downv/ash previously established.
The limit CL boundary shown in Figure 3-72 represents the limit oflinearity in the lift-curve slope data. These values of CL correspondclosely with an abrupt positive break in the pitching moment data. Thesteepness of this boundary at the design Mach number of 0. 98 indicatesthe need for additional wind tunnel data in the transonic and low super-sonic Mach range. Based on the assumptions of the analysis, adequatetrim power is available for the useable range of lift coefficients. Enginethrust effects v/ere not included in the trim equations.
An indication of maneuver capability is shown in Figure 3-73. Normal.'load factor is shown versus altitude for a gross weight range of 1SOO to2400 pounds. The lift coefficient at each Mach number corresponds tothe limit CL boundary of Figure 3-73.
Conclusions
Based on the results of this preliminary analysis, the existing BQM-34Ehorizontal tail appears adequate for both longitudinal stability and trim,in conjunction with the supercritical wing configuration designated asconfiguration 1-30-2. Actual downwash data should be available formore refined analyses, and inclusion of aeroelastic effects should-beconsidered if flight at high dynamic pressures is envisioned.
NASA wind tunnel data indicate nonlinear pitching moments at a high liftcoefficient. A better definit ion of this characteristic, by means of windtunnel tests of a scale model of the research drone, would be desirablein the transonic Mach number range.
144
—1
Figure 3-72. Model BQM-34I-; W i t h Supercr i t ical Wing' 1-30,Elevator An^le Required tor Trim
"DO
'Si
OJO
I
ci>
Og,
GO
CO
C5"CQ
TJOs.
COc-
Icoas-,
146
Fore-Body Effects
An estimate has buc-n made, by means of Out com methods, of the effect oflengthening the fuselage on the static longi tudinal stability of the subjectresearch configuration.
Lift and pitching moment characteristics of the nose and forebody arcdiscussed in Paragraph •!.;}. 2.1 of Datcom, and the following equationsare given to determine the increments of the lift and pitching momentcurve slopes:
2(k -k ) S2 1 oCL = ^~ , per radianv
2(k2"
kl) (' dSxC = / — — (2, - x) dx, per radian
m V, / d ba. b. •' x
o
The nose and forebody are considered as the fuselage section forward ofthe wing-fuselage juncture. The equations above, evaluated for a body-length of 12.6 feet, give the following:
C = 0.00321 per degree, based on wing geometry.L
O.
(j = 0. 00491 per degree, based on wing geometry, relativeO. to base of forebody.
As a check on the validity of the method, the equations were also used tocalculate the stability level of the entire fuselage,' and a comparison wasmade with wind tunnel test data:
DATCOM (REFERENCE 1) TEST DATA (REFERENCE 2)
C per deg. 0.0023 0.00250!
C per deg. 0.0112 0.0130 M < 0.40ma.
Based on the data from either source, the effective center of pressure ofthe fuselage is about one MAC forward of the nose, indicative of thecouple produced by bodies in potential flow. If the moment due to thenose lift is doubled to approximate a couple and a viscous cross-flowterm is added, the resulting moment curve slope is on the order of 0.0105,which is in fair a r e e m e n t with the above values.
The effect of additional fuselage length was determined by adding constantarea sections forward of the base of the forebody and determining theincrement in lift and moment due to the additional volume.
C x Vm b
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The moments were then referenced to the center of gravity, and thoincrement in static stability was determined from
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The m a x i m u m dynamic -pressure was determined from the speed-alti tudeand corresponds to a Mach number of 1.36 at 10,000 feet.
3.3.4.2 Lateral -Di-reetional Stabil i ty
A brief analysis lias been conducted to evaluate the lateral -directionalstatic stability of the EQM-34E equipped with the supercritical wingdesignated conf igura t ion 1-30-2. The results indicate that the BQM-34Evertical tail will provide positive direct ional stability/ but that it may bemarginal at low Mach numbers . Ut i l iza t ion of the directional s tabi l i tyaugmentation system may be desirable for satisfactory stabi l i ty charac-teristics at low speeds and high angles of attack.
Dihedral Effect
o estimated for conf igura t ion 1-30-2 by subtracting the estimatedvertical tail contribution from NASA \vind tunnel model data and addingthe contribution of the BQM-34E vertical tail. The change in C % 0 dueto the low-wing location was es t imatedby an empirical expression in E tk in , p. 4SG. The resulting level of C% p
is shown in Figure 3-75,
Directional Stabilitv
Comparison of the vertical tail geometry of the NASA model and the BQM-34K shows approximately the same ta i l vo lume coeff ic ient , V = 0.13,
E
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Figure 3-74. Model BQM-.14E E-stiniated Cliangc in Stabi l i ty due toAdded Fuselage Section Forward of Win"'
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for the same ratio of exposed area to ju'oss-tail area as established by theBQiM-;!-H:.'. The NASA model vor t ica l t a i l ' j imlr ibut ion to Cno could notbe determined accurately in the absence of. tail-off data, but it is esti-mated to be approximately Ifj percent more effective than the BQM-3-1Evertical due to a higher aspect ratio and lo\vcr sweep angle. Thisresults in a reduction in Cn „ 01 0.00078 for the subject conf igura t ion , asshown in Figure 3-75. Also shown in the figure is the ratio of C^ fi toCng » which is an indicator of dutch roll characteristics. This ratio isabout the same for the subject configurations in the design Much numberrange. However, at low Mach numbers corresponding to launch airspeeds,the ratio increases and approaches that for the basic BQM-34E with theexternal lank on. Experience with air launch of the BQM-34E indicateda need for high directional stability. The directional stability augmenta-tion system is used for this purpose for the tank-on configuration andmay be desirable for the research configuration utilizing a standardBQM-34E vertical tail. The directional stability can be increased byapproximately 0.001G and will increase Cno to a level equal to or higherthan the NASA model data. Other factors would influence the closed-loopand dynamic stability characteristics, such as'yaw clue to roll control,higher roll and yaw inertias, and higher roll damping from the super-critical wing.
The effects of angle of attack were not checked, but the increase in C^ ,,Pwith a. for the NASA configuration indicates that a higher level of Cnnthan that provided by the basic BQM-34E vertical tail may be desirable.
3.3.5 Mission Performance
The performance capabilities of configuration 1-30-2 in the Mach-altitudeplane was included in the point design summary of Paragraph 3.2, Figures3-17 and 3-22. In addition to this, an actual time history of an examplemission at maximum power was made for both a ground launch and an airlaunch at 10,000 feet. These results, shown in Figure 3-76, indicatethat this vehicle should be able to provide on-station mission data atspeeds close to Mach 0. 93 for over 20 minutes. Additional performancecapabilities are illustrated in Figures 3-77 and 3-78.
r
3.'3. 6 Command and Controli
The preceding discussion has been concerned mainly with the feasibilityof modifying the BQM-34E airframe for research applications. This sec-tion discusses the other vehicle subsystems, primarily avionics, whichmust also be modified. These subsystems include the data transmissionlinks, automatic flight control, and secondary power (which is not
" ' 'I II-~ J~ ' it. ILu
Figure 3-76. No. I--JA oJ°-
Figure 3-77. No. 1-30 Maximum Ratc-of-Climb vs. Alt i tude
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"commonly included under the avionics label). The propulsion andrecovery subsystems remain unal tered. Analysis of the ground-basedportion of the vehicle control system was beyond the scope of the study.However, the influence of (he availabili ty and performance of groundequipment on the airborne equipment had to be considered for complete-ness.
A simplified functional block diagram of the'airborne subsystems (lessairframc) is shown in Figure :;-79. Of the 12 functional blocks shown,three are new (for wing controls) and one is modified (automatic flightcontrol system (AFCS)); the remainder are unchanged from the targetconfiguration with one q u a l i f i c a t i o n . The recommended command guid-ance transponder has not yet flown in the BQM-34E/F; however, it hasflown in several versions of the DQM-34A subsonic drone. The functionalmodifications for installing the transponder amount to interfacing with theAFCS; hence, the modif icat ion is allocated to the AFCS.
The difference between the E (Navy) and F (Air Force) models of BQM-34Elies in the target augmentation equipment complement. Hence, withthese equipments removed, the models are virtually identical. :
The target command and control system is designed to be operated bymilitary personnel having a minimum of training. The operators aregenerally not pilots. The controls available to the operator are discrete(r. e., relay closure) commands, limited flight data for performance mon-itoring, and a vehicle tracking display. The commands are limited tosuch as TURN RIGHT, T U R N LEFT, CLIMB, DIVE, turning equipment .on or off, initiation of the recovery mode, etc. Each maneuver commandenergizes a potentiometer in the drone autopilot, which is set prior toflight to a specific command voltage. The autopilot responds to the com-mands in a proportional manner. For example, the autopilot responds toa turn command with a constant-altitude turn whose roll angle (and con-sequently turn rate and load factor) is proportional to the preset voltage.
Once initiated, the turn relay remains latched until it is disabled.by aSTRAIGHT ANt) LEVEL command. The latter command represents -/.eroroll angle to the autopilot. The turn command potentiometer voltage can,therefore, represent any roll angle from 0 through about S5 degrees, aslimited by maximum load factor, although only one value can normally becommanded during a flight. (In special modes, an alternate level can beselected.) The autopilot responds to CLIMB and DIVE commands, in asimilar manner, as altitude change commands. Mach number can alsobe controlled, although it is normally not commanded directly by tl\eoperator. lie docs so by keeping engine rpm (i.e., a discrete command
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of throttle rate) whi le monitoring engine rpm and Mach number 0:1 flightdata readout. Oilier displayed flight data parameters include pitch androll att i tude, a l t i tude , and heading. For a complete discussion of thetarget control system, see References 17 and 18.
A proportional command capability is achieved by replacing the discretecommand l ink with one having continuously variable data channels, as intelemetry data, and int roducing the command variables into the autopilotin lieu of potentiomet'^^feltages. Hence, the autopilot responds to acontinuously variabf^^^^ge command that is controllable from theground.
Avionics Reconfiguration
The. avionics: (including secondary power-and servoactuators) set requiredfor research operations are derived from the basic BQM-34E/F target javionics in the following manner: |
II
a. The target augmentation and scoring equipments are removed.These include such items as radar augmentors and antennas,infrared sources, miss distance sensors, and tow-targetequipments. ' :
b. The. standard command receiver and telemetry transmitter jare replaced'by a command guidance transponder and possibly ja small, wide-band telemetry transmitter. j
c. An electrically driven hydraulic power supply is added for jthe wing control surface actuators. A small electronic unithousing the flutter mode control computer and ancillaryAFCS interface is added. '•
d. The resulting equipment complement is rearranged withinthe compartment to utilize the available space to better
j advantage. The cooling system was retained some.whatI arbitrarily since it is not required, except at the highestI Mach numbers. For many research operations, the cooling! system space can be occupied by other equipment.I
Command Guidance Data Links
The command guidance data links provide the means for communicatingwith the vehicle for purposes of control and remote measurement. Theimportant link parameter is its frequency bandwidth'or channel capacity,
.which denotes the amount of data that the link can transmit. The factorsthat determine the bandsvidlh required are the number o[ parameters tobe transmitted (i.e. , the number of channels), their resolution or accuracy,and their frequency content, which is also bandwidth.
Multiple channels arc obtained either by dividing the available bandwidthinto narrower bands (frequency-division multiplexing') and transmit t ing allparameters simultaneously or by transmitting samples of each parametersequentially ( t ime-divis ion mult iplexing) or a combination of the two (sub-multiplexing). Most available drone control systems use time-divisionmultiplexing. The important factor in time-division multiplexing (TDM)becomes the sampling or update, which must be at least twice as high asthe bandwidth of interest of the parameter to be transmitted.
The bandwidth and sampling rate required of the drone data links dependon the guidance and control philosophy employed.' The philosophy deter-mines-which control laws are mechanized and whether control loops areclosed in the air or on the ground {whether by man or computer). Suchalternatives are indicated in Figure 3-80. In the figure, the width of thedata link arrows is proportional to the bandwidth required. The basicvehicle control loops are indicated in the four blocks with their relativefrequency ranges indicated. The AFCS outer loops include airspeed,altitude, and heading control and -phugoid modes. These control para-meters are used in a typical target. The difference between the top twoblocks is proportional command versus preset discrete commands. AFCSinner loops include short-period dynamics, stability augmentation, andhandling qualities. This is the loop in which a man operates in an aircraftwithout an autopilot. Note that the inner loop control frequencies are anorder of. magnitude greater than the outer loop frequencies. Wing fluttermode control loops (and body bending as well) involve frequencies thatare another order of magnitude greater than those of the inner loops.
The sampling rate on the telemetry must be somewhat greater when theflight control computations are performed on the ground rather than in theair to prevent the overall transport delays in the closed loop from dis-toring the response or causing oscillations. . The increase would be on theorder of a factor of two to four times as great. Similarly, the quantizationlevel (digital resolution) is important, because too large an increment cancause limit cycling. For example, if roll attitude were to be quantizedover ISO degrees with an S-bit word (256 increments), then the resolutionwould be 0.7 degree. This could produce a small-amplitude limit cycleof a similar magnitude. Hence, a longer word, say 10 bits, would bedesirable.
159
CONTROLCENTER
CONTROLCONSOLE
I• COMPUTER -
. COMPUTER
COMPUTER /
COMMAND/TELEMETRYDATA LINK
VEHICLE
PRESETMANEUVERS
AFCSOUTER LOOPS
AFCSINNER LOOPS
FLUTTERCONTROL
LOOPS
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"= CONTROLFREQUENCIES
(RADIAUS/SECOND)
Figure 3-80. Guidance and Control Concepts
160
The command guidance equipments most o.Cten used today for drone controlarc listed in Table :'}-'21. Kuch ui tiie.se equipments is capable o f h a n d l i n gthe research application wi th respect to conn-oiling vehicle maneuverdynamics for either ground or a i rborne computation. However, withrespect to the capability of h a n d l i n g wing-flutter and body-bending modecontrol, each would be l imi ted to f irs t modes at best, particularly withground computation, because the closed-loop delay times become signifi-cant at such frequencies. Possible solutions to the latter problem are asfollows:
a. Develop or adopt a new or wider bandwidth telemetry l ink.
b. Modify the exist ing links to increase their bandwidths.
c. Compute only in the vehicle and transmit the flutter datato the ground for monitoring purposes over a separate dedi-cated (standard) telemetry link.
The options are listed in order of decreasing cost, schedule impact, andflexibility.
The main differences between the equipments listed in Table 3-21 are inthe tracking function and-relative cost. The Vega system uses the localtracking radar as a host for its telemetry carrier, the Motorola systemhas an integral tracking radar (with lower power and accuracy), and theBabcock telemetry link is separate from the local tracking radar. TheMotoi'ola system also has an integral control and display console, whichthe other two do not. The equipments are listed in order of generallyincreasing cost, although the f i rs t two are significantly lower thanMotorola because of the ;r lesser complexity.
Automatic Flight Control System
The BQM-34E/F AFCS provides control of vehicle altitude, Maeh number,pitch, and roll attitude plus three-axis stability"augmentation. A detaileddescription of the AFCS function and performance can be gained fromReferences 17 through 19.
Since the target AFCS was not designed for research applications, itlacks some of the flexibility and capability which are desirable. However,for the application being considered, it can satisfy the immediate require-ments with some relatively simple modifications.
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The options available are as follows:f
a. Retain the AFCS as i.s, except for modifications forproportional command inputs.
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b. Augment the option 1 AFCS with the necessary computationand logic for new modes.
/c. Retain the AFCS for launch and recovery, but bypass the
sensor and/or computer sections.with airborne or ground-based alternate equipment for the test portion of flight.
d. Replace the exis t ing AFCS with a new one having capabilitiesmore suitable to the application.
In general,..the options are listed in order of i-ncreasing capability, cost,and development time. Tcledvuc Ryan has successfully flown aircraftemploying options a, b, and d. NASA Flight Research Center (EdwardsAFB) is about to fly a spin test vehicle using option c without AFCS.
The recommended course of action is option b. The sorts of modifications,required within the AFCS are outlined as follows. The longitudinal andlateral axes are sufficiently separable functionally to be consideredindividual!}'. A simplified block diagram of the longitudinal axis, includingrepresentative modifications, is shown in Figure 3-S1. The sensors cur-rently used (air data, vertical gyro,, rate gyro, and normal accelerometer)are those which would be expected in a research vehicle.
Existing command inputs (continuously variable) include Mach number,altitude, and attitude. Rate or acceleration command mode can beobtained by introducing switching "logic prior to the stability augmentationsumming junction. The command would be shaped prior to summation toprovide the proper response characteristics, in the manner of commandaugmentation. In operation, the al t i tude, Mach, and attitude Inputs wouldbe diverted to the synchronize mode, so that reversion to one of those•modes would.not cause a switching transient.
iiThe figure also shows aileron servo inputs, to indicate how collectiveailerons or flaps could be driven for direct-lift control studies.
iThe lateral axis (modified) is depicted in Figure 3-82. It consists of theyaw and roll axes plus the flutter mode control subsystem. The yaw axisis shown as it currently exists with two exceptions: provisions for yawcommand are included, and the sideslip sensor is used for contrpl only
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when the external belly fuel pod is attached. The roll axis is also shownas it currently exists with two exceptions: provisions for roll rate com-mand are included, and the ou tpu t drives aileron servos rather than theexisting rolling tail sei-vos. The tail would then operate purely as anelevator.
The f lut ter mode control loop consists of a number of wing-mounted sen-sors (accelerqmetcrs and possible rate gyros), a compensation filternetwork (approximately f i f t h order), and \ving-surface servoactuators.The filter mechanization in.analog form is straightforward. It would con-sist of several operational ampl i f ie rs and a resistor-capacitor pair foreach filter clement. 'The filter characteristics (i.e. , gain and time con-stants) can be made variable and can be controlled remotely if needed.For example, each axis of the stability augmentation system has avariable-gain element which. is controlled by a voltage. The controlparameter is dynamic pressure,, but a remote command, being a voltage,could also control it.. Because the flutter mode frequencies ef interest •approach 30 Hz, which could overtax the existing data links, and becausethe filter design is straightforward, the onboard mechanization is recom-mended.
Secondary Power
The BQM-34E/F secondary power i-s derived from an engine-driven dcgenerator, which also serves as the engine start motor. Hydraulicpower for the control-surface actuators is provided by an electricallydriven supply. The servoactuators and power supply form an integral ,self-contained unit, which is also a structural member of the air-frame.
Additional hydraulic pov. 3r will be" required for the wing actuators. Theexisting supply is sized for the tail actuation requirement: hence, it doesnot have any significant reserve capacity. The engine has only one powertakeoff pad, which is used by the generator. Therefore, since an engine-driven hydraulic pump is not possible, the alternative is to provide anelectrically driven hydraulic supply. A number of such supplies are usedoh missile and reentry vehicles. They are small enough to fit easily intothe drone. Further, sufficient electrical power is available to drive one.
iThe hydraulic power requirements have been estimated as follows: themaximum aileron hinge moments for the inner and outer ailerons andleading-edge flaps are 1500, 1000, and 500 inch-pounds, respectively;the frequency response of the outer aiteron/leading-edge flap pair shouldbe at least 100 radians per second at the first order break; and the inneraileron response should be about 20 radians per second. The existiu«v
hydraulic system supplies two actuators having 4000 inch-pounds ofstall torque and one (rudder) having 900 inch-pounds of torque stall .All servos have a f i rs t -order lag of slightly greater than 20 radians persecond.
A very gross comparison ot" power requirements can be obtained by multi-plying stall torque by fn.-quency response for each servoactuator andsumming. When this i.s done for the wing set and tail set, the ratio ofwing power/tail power is approximately two.
The electrical' input to the tail hydraulic pump is 20 amperes at 28 voltsdc. It supplies O.G gallon per minute at 1000 psi. Data on two availableelectrically driven power supplies is presented below. Note that thisdata indicates that the power supplies can provide 2-1/2 times the powerof the existing supply.
' TYPE
Pesco Model165-100
Pesco Model144-300
USED ON
Martin hypersoniclifting body
Minuteman ThirdStage
PRESSURE
1500 psi
1500 psi
FLOW
1 . 0 gpm
1.0 gpm
ELEC-TRICALINPUT
28 Vdc38 amp.
28 Vdc38 amp.
The available electrical power, summarized in Table 3-22, is adequatefor driving either supply while retaining a reserve for additional equip-ment.
Conclusions
The feasibility of converting the BQM-34E/F avionics from a target-oriented to a research-oriented configuration has been analyzed with thefollowing results: _„
a. The modifications are confined primarily to the automaticflight control system arid to equipment relocation.
b. Command guidance data links with adequate capacities forresearch applications are available.
167
TABLE 3-22
S K C O N D AH Y PO W E R
©AVAILABLE GENERATOR CAPACITY 200A @ 2S VDC
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WARMUP 134A
CRUISE 94A
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RESIDUAL CAPACITY : .
WARMUP . 29-11A
CRUISE. . . . ; . . . . 6G-4SA
168
3.4
c. The control laws are well understood, and their mechaniza-tions arc u - i t l i i n the current state of the design art andhardware capability,
d. Adequate electrical and hydraulic power are available.
SUPPORT STUDIES
3.4.1 Wind Tunnel Tests
To assure a high probability of success of new or revised RPVs, it isrecommended that aerodynamic lest data be obtained in each of the criti-cal flight regimes. This will provide not only a verification of estimatedaerodynamic, stability, and control coefficients but, in addition, willmake possible realistic preflight simulations, including nonlinear effectsdue to compressibility and separation phenomena.
For the subject vehicle, this would include low-speed transonic as wellas supersonic wind tunnel test data of scale models, as required, closeto flight Reynolds numbers. Although new vehicle checkouts usuallyinclude engine inlet tests, boundary-layer gutter optimization, etc. , it,is felt that this is not likely to be required for the subject application.The basic inlet configuration is designed to operate, with reasonablecompromise, in both the subsonic as well as the supersonic regime.(Mach 2.0 tests indicated a mild instability.)
A requirement for pressure taps to provide good chordwise and spanwiseload data is always desirable, from an analytical viewpoint, in both aero-dynamic and load analyses. However, this requirement is seldom imple-mented, because of economic and time constraints.
Typical flight modes of a new wing to be critically examined by means ofwind tunnel tests would include the following:
a. Low-speed launch mode, Mach 0.1 to 0.4, free-fall stabilityand trim at near zero lift.
i
I b. High-speed launch mode, Mach 0.6 to 0.8 (only if required).i
i c. Maximum climb trim and stability, Mach 0.4 to Mniax.
d. Cruise trim, stability, and control, Mach 0.4 to Mmax.
e. Maximum load factor (turn mode).
f . Po\ver-off glide characteristics, M a c h O . S t o 0.2.
g. Recovery 1111>de, drag chute.
h. Maximum t r i m C[, versus Mach number.
i. C a p t i v e - f l i g h t leads on carrier aircraft./
•3 .4 .2 Fl ight Assurance Summary
Reliability
Flight-phase and recovery-phase inherent reliability predictions for theNASA-configured .BQM-.'M l\. have been completed. These predictions weredeveloped from BQM-:ML' reliabil i ty prediction mathematical models,with adjustments for the currently planned changes to the Navy vehicle.Sixty-five minutes (1.053 hours) ' f l ight phase, and 22 minutes (0.363 hour)recovery phase durat ions (Navy prediction profiles) were used to providea comparison of the two vehicles. The maximum phase durations wereselected to provide a conservative estimate of inherent reliability. The .results are as follows:
NASA NAVYBQM-34E BQM-34E
Flight phase • 97.909£ 98.04%(With cooling system installed) 97.79% 97.93%.
'Recovery phase 99.42% 99.63%Recovery and retrieval (combined) 98.0 %
These values are for the- air vehicle shown in Figures 3-83 through 3-84.
The flight phase includes the period from launch to the initiation ofrecovery procedures. The recovery phase includes the period from theinitiation of recovery procedures until the air vehicle is in a position tostait the retrieval operation. For this analysis, the worst-case conditionof parachute descent to a water landing was assumed.
i
Since there is no reliability model for a Navy MARS retrieval system,data from other programs in which the MARS system is used was examined.A combined recovei-y and retrieval reliability of 98.0 percent is indicated.
The NASA BQM-34E predictions are based on the system changes discussedin the following subparagraphs.
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Airframo. - The NASA-WO wing is substituted for the Navy wing. It isassumed that the fa i lure rate is equal to four times the Navy wing fai lurerate due to added complexity and planned fl ights approaching the wingstructural l i m i t s . Planned f l igh t s beyond the structural l i m i t s have beenexcluded from this analysis and will require further study during thedesign phase . ) Four wing trailing-cdge control surfaces, each with afailure rate equal to one horizontal stabilizer, arc added, as well as twowing lead ing-edge control surfaces, each with a failure rate equal to one •rudder.
^Propulsion System. - No change is made in the propulsion system.
.Electrical System. - The Air Force power distribution box failure rateis substituted for tha t of the Navy power distribution box to provide forpotential increased funct ional requirements.
Flight Controls. - An elcctrohydraulic actuator.with a failure rate equalto those of existing electrohydraulic actuators is added for wing controlsurfaces. Additional flight control box functions, with a combined failurerate equal to the combined failure rate of the existing pitch commandassembly and 0.5 times the existing relay logic assembly, is provided.
Guidance, Telemetry, Tracking. - The existing radio receiver andtelemetry transmitter are replaced by the (Vega) VTCS, and an a. sensorwith a failure rate equal to that of the existing £ sensor is added.
Equipment Cooling. - This system is not currently planned for use;however, air vehicle reliability is shown for both cases (i.e. , withoutor with the cooling system installed) in the event that supersonic flights.may later require the system be installed.
Recovery. - The Air Force MARS main parachute system is sub stiltedfor the Navy main parachute system. •
Table 3-23 shows the NASA BQM-34E and Navy BQM-34E flight phasereliability prediction comparison on a system by.system basis. Table3-24 shows the same comparison for the recovery phase.
Maintainability
A preventive-maintenance man-hour analysis for the NASA BQM-34Ewas performed based on the Navy BQM-34E maintenance engineering
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analysis report. The results arc compared with the predicted anddemonstrated preventive maintenance man-hours for the Navy BGJ1\I-34E,which docs not have the "A1A1JS system, as follows:
NASA BQM-34E NAVY BQM-34E_J\vith MARS) (no MARS)
Estimated 114.90PMMH 1GG.19PMMH
Demonstrated — 178.37 PMMH
These are the direct, average, preventive-maintenance man-hours perflight. The NASA estimate i.s based on the following assumptions:
a. This estimate is for the second and subsequent .flights. Thefirst fl ight requires an additional 12'man-hours if uncratingis considered.
b. The flight control system will require 75 percent additionalman-hours due to additional flight control system functions.
c. MARS recoveiy is used.
d. Maintenance man-hours are direct (i.e., "screwdriver-,time") man-hours.
" e. Maintenance hours do not include time for operational taskssuch as uploading, prelaunch tests, launching, flight, orretrieval;
f. The cooling, system is not used.
g. Augmentation (for target missions) is not installed.
h. Test time for the VTCS (Vega system) is equivalent to thatfor the AN/DRW-29 receiver and the AN/AKT-21 TLMtransmitter.
i. A ground launch is assumed.
Table 3-25 shows the breakdown of the separate task estimates.
TABLE 3-25
KSTIMATl- OF TASKS
No. .
1.
2.
3.
4.
5.
6.
• 7.
. 8.
Q _
10.
11.
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16.
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Task D e s c r i p t i o n
Systems C o n f i d e n c e Test 1(Completed V e h i c l e )
Service Vehicle with Fuel
Weigh Vehic le
Assemble t ; A l i p p . RATOBottle to A t t a c h F i t t i ng
Service Bat te ry
Preflioht Servicing
Disassembly a f t e r F l i g h t(Remove Equip. Connp.Doors, ADC, G yr os , e t c . )
Check Components
Pressure Checks
Prepare for Instal ledEngine Run
Installed Engine Run
Prepare for SystemsTests ,
Install Equipment inEquipment Compartir.er.t ••.-
Perform Systems Tests
Complete Assembly ofVehicle
• B u i l d up and I n s t a l lRecovery Svstem (In-cludes MARS)
Weigh and BalanceVehicle
TOTAL PMMH
Clockl i o u r s
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0. 50
0. 50
1. 50
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1. 90
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3. 00
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1. 90
1. 20
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2. 00
3. 00
8. 50
8. 50
36.60
11. 50
16.60
6.00
114.90
1 7O
Component Test Requirements
The current NASA BQM-.TIK configuration will require only one newmajor component that wi l l not have demonstrated flightworthincss. Thisis the wing control surface actuator package. Assuming it is-a unit com-parable to the existing eleclrohydraulic actuator, it is recommended thateach uni t procured be subjected to a flight-assurance test equivalent tothe reliability sampl ing test performed on the selected units procured forthe BQM-3-1E. The test profile includes low and high-temperature soak,low and high- temperature operation, three-axis vibration, an acceptancetest, and visual inspect ion. After successful completion of this test,each unit will then be refurbished for flight readiness .and subjected tothe acceptance test procedure prior to shipment from the supplier.
Elements to be considered in a flight-assurance determination arepresented in Table 3-2G.
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4.0 CONCLUSIONS
As a result of this feas ib i l i ty study, it is concluded that the basic BQM-34E is.readily amenable to modificat ion for conversion to a NASA I'csearehdrone.' Wings were sized to indica te the applicability of the BQM-34E toa wide range of subsonic and supersonic missions. Six point designs withresearch \vings applicable to advanced transports, RPVs as well as anair-to-air fighter, were ident i f ied . Comprehensive structural, and designanalyses \verc accomplished on a representative research configurationto indicate practical modif ica t ions to provide high and low-wing structuralattachment capabilities. Typical inboard and outboard ailerons and activecontrol devices were configured with practical actuation system arrange-ments/ Cost-effective methods of constructing wings with various degreesof bending and torsional rigidity were determined, for possible loads andflutter suppression research studies.
The required modifications to the existing command and control system,to provide capabilities of accomplishing control law functions via ground-
', based or'airborne computers, were identified within the state of the artand available avionic systems.
"Page missing from available version"
5.0 R ECOM M E NDAT IONS
According to the results of this study, the basic BQM-34E drone systemis readily adaptable into an unique NASA free-flight research systemcapable of accomplishing both subsonic and supersonic tasks. ROM costs,delivered to the customer (per Reference 1G), indicate that this researchdrone can provide substantial savings in terms of time and resources inthe development of man-rated systems. Free-flight validations withouttunnel-wall constraints can readily be established in critical flightregimes and where wind tunnel test data are in question (such as at Mach1.0). It is therefore recommended that such a program be pursuedimmediately to provide NASA with this capability within, time schedulesindicated in Figure 5-1.
185
CON£lGUP. A7 in:.' CM" i 1MT1ON
• r-:CSI2l.'.'G AN1;.) MATCHING ANALYSIS
• CONF tGuiw\7;oN Tr-.. '-r<rs«. co .•-.' f K; u •( A T i o .'.• f i •<• i: n z c
\Vii'>J_[j TUMNTl. 1ST.7S
o rAiinlCA"! ION' AND C H E C K O U T
o WiN'D TUN.'.'t. L 1 i S ' fS AN'O HAT A f'i.'.r'ORI
o WiNG AND HIGH L i T T DhVlCl'S
o ruscLAc.c .v.o;./. F ;V;AT;O>;
o CON'1'IU.n. A.'vi.) CU'.'.'.'AWD SYSTCM
* WING Af.T) CON'Ti'iOLS
» fUSf lLAGL
, AViOMCS
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G.O NOTATIONS AND SYMBOLS
Conventional notations are used throughout this report. They arc listedas follows:
A Wing aspect ratio
AVSYN Air vehicle synthesis program (Teleclyne Ryan)
b Wing span
c Wing mean aerodynamic: chord
CFE Equivalent f la t -p la te drag coefficient
C Wing-root chordr
C Coefficient of friction
C Wingtip chord 'V
eg Center of gravity
_, „ r,. . , DragC Drag coefueient, ——•D qb
Base DragC Base drag coefficient ,
D, b qb
G Drag coefficient at zei'o lifto
C2~ Drag-due-to-lift parameter
L
C ' Lift coefficient,L qSI
C Slope of lift curve, per degreeL
^ T,., , . „ Pitching MomentC • Pitching moment coefficient, °
m qS—c
187
C /C Longitudinal - s tabil i ty parameterm L
C /Sir Pitch control effect iveness of horizontal tailin H
s
C Directional - s tab i l i ty parameter, per degree
fi
g Acceleration clue to gravity
h • Altitude, feet
K ( Relative engine size to base reference engine lift-to-dragN ratio
L/D Lift-to-drag ratio
M Free-stream Mach number
NX Normal load factor
q Free-stream dynamic pressure, pounds per square foot
R Distance, nautical miles!
RN Reynolds number
rpm Revolutions per minute
5 Reference whig area, square feet
TOS Time on station, minutes
_ 'Si SuV Horizontal tail volume coefficient. = -^r x
H e Sw
— ^v SvV Vertical tail volume coefficient, = —- x
v b Sw
W/S Wing loading, psf
WT Weight, pounds
Wb Body width, feetAngle of attack, degreesAngle of sideslip, degrees
6 Horizontal-tail, deflection, degreesn
C Effective downwash angle, degrees
F Dihedral angle, degrees
• 188
SUBSCRIPTS
nia-\ M a x i m u m
B Body
c Cruise
H Hori/.oiUal tail
V ' Ver t ica l tail
W Wing or \vetted area
REF Refer once
0 Zero l i f t
180
"Page missing from available version"
\qo
T.O REFERENCES
NASA Memo 33VpleasairtKj 20
Contract MAS1-11758" .1972, "Stxidy Guidelines for
5
6 ,
7. '
8.
9.
10.
11.
12.
13.
14.i
•15 •!
l6.
17.
Proposal for a Feasibility Study of Incorporating a ResearchWing on a BSU-o ITProne, ASTi-I 72-13, 12 May 1972.
Task I Results. NASA Wing Feasibility Study EQH-3te, ASTM72-22, 30 August 1972.
NASA Letter 126A/WA SI -11758, dated 30 October 1972.
Statement of Work, Fear, ibility Study of Incorporating ResearchWing on £014-3 Drone, l-lb-2662, dated 3 March 1972 /
Basic Data Report for BQM-3k-E Supersonic Aerial Target,Report Ho. TEA Ibb2671s7 25 May 1970.
Actual Weight Report for BQI-I--3 E Supersonic Aerial_ Target,Serial Nuniber J>'/-l6lcJlv Report Ho. TRA 166W-22, dated4 May 1972.
_. Re-oort for EQM-3^F Supersonic Aerial Target ,Report I-Io. "TEA' loo5'l-25 _, "dated 4' February 1972.
NASTRAN Prograa, Office of Technology Utilisation, NASA,1970.
Bo-wers, N.M., Coapendiuia of Computer Data from HASTRAN.
Stress Memo 8Sa, LockJieed Aircraft Corp., 15 September1955..
"Buckling of. Plat Plates", Handbook of Structural Stability,Part-1, IIACA TH-3781.
Peery, D.M., Aircraft Structures, McGraw-Hill Book Co., Inc.,1950.
Letter of Transmittal 7/60/3826, Santos to NASA ContractingOfficer, ROM Costs, dated Ik December 1972.
Brown, W . R . , Guidance System Ji§i|£,t for BQj--l-3 E Supersonic^Aerial Target, Report iio. l6bloS;, 19 March 1970.
Downey, P.J., Control and Stabi].ization System Report forEOJ-i-3^E Supersonic Aerial Tax-get, Report Ho. TRA 16654-4E,12 June 1972.
Wooley, M., and Jones, J.G., System Dynamics Report forBQH-S^F Supersonic Aerial Target, Report No. TRA' 15654-3D ,Supplement No. 1, 20 January 1971.
191
