Airloads research study. Volume 2: Airload coefficients derived from wind tunnel data

CONTENTS Page Section SUMMARY nITRODUCTI ON AIRCRAFT DESCRIPTION Fuselage Wing Nacelle Horizontal Stabilizer Vertical Stabilizer Stability and Control Augmentation System STRUCTURAL LOADS DATA DERIVED FROM WIND TUNNEL TESTS METHODS AND RESULTS REFERENCES APPENDIX A. SYMBOLS APPENDIX B. RIGID AIRLOAD COEFFICIENTS AT LOAD REFERENCE POINTS - M=O.85, A =67.5° W APPENDIX C. NET RIGID AIRLOAD COEFFICIENTS FOR FLIGHT CONDITIONS APPENDIX D. FIGURES USING ENGINEERING UNITS APPENDIX E. RIGID AIRLOAD COEFFICIENTS AT LOAD REFERENCE POINTS - M=1.20, Avf67.5° APPENDIX F. RIGID AIRLOAD COEFFICIENTS AT LOAD REFEREN,CE POINTS - M==O.95, A =67.5° W Rev Oct 1981

CONTENTS (Concluded) Section Page APPENDIX G RIGID AIRLOAD COEFFICIENTS AT LOAD 75 REFERENCE POINTS - M=0.70, Awr67.So APPENDIX H RIGID AIRLOAD COEFFICIENTS AT LOAD 84 o REFERENCE POINTS - M-1.60, A =67.S W APPENDIX J RIGID AIRLOAD COEFFICIENTS AT LOAD REFERENCE POINTS - M= 2 . 20, ~=6 7. 5 ° Added Oct 1981 iv LIST OF ILLUSTRATIONS Figure Title Page 1 B-1 aircraft ....

2 Structural breakdown :5 Wind tunnel pressure loads model 4 Unit additional span load distribution on wing - centerbody AW = 67.5°, M = 0.85 ...... .

.5 Unit additional chordwise load distribution on wing' - centerbody AW = 67.5°, M = 0.85 6 Normalized pressure distribution . . . . . . .

7 Flight control surfaces ........... .

B-1 Component rigid airload coefficient reference points, AW = 67.5° . . . . . . . . . .

D-l General arrangement - RDT&E A/C-l and -2 D-2 Fuselage structure diagram (B-1) (-55B).

D-3 Wing-structure diagram, outer (-55B) ..

D-4 Structural arrangement - nacelle external compression inlet (RDT&E). . " . . . . . . . . . . . . . . . . 51 D-5 Horizontal stabilizer - STRL arrangement (42-1/2° LE sweep) ................ .

D-6 Vertical stabilizer structural arrangement . . . . .

Rev Oct 1981 v LIST OF TABLES Table Title Page I Aerodynamic Effects Applicable to Component Loads. 14 B-1 Wing Coefficients at X 354 for 0.85M and RS I\.W = 67.5° . . 27 B-II Horizontal Tail Coefficients at BP 10.75 for 0.85M and I\.W = 67. 5 ° . .

B-III Vertical Tail Coefficients at WL 136.56 for 0.85M and I\.W = 67.5°. . 29 B-IV Forward Fuselage Coefficients at FS 528.5 for 0.85M and I\.W = 67.5° . .

B-V Aft Fuselage Coefficients at FS 1337.5 for 0.85M and I\.W = 67. 5 0. • 31 Vertical Tail Coefficients at WL 75.0 for 0.85M and B-VI I\.W = 67.5° . . . . .

B-VII 33 Load Coefficients Data Card Listing - M=0.85 Wing Coefficients at XRS 354 for 1.20M and E-.I ,~ = 67.5° E-II Horizontal Tail Coefficients at BP 10.75 for 1.20M and .'\W = 67.5°.. . . .. .

E-III Vertical Tail Coefficients at WL 136.56 for 1.20M and AW = 67. S° . ....

E- IV Forward Fuselage Coefficients at FS 528.5 for 1.20M and AW = 67.5° . .. .

E-V Aft Fuselage Coefficients at FS 1337.5 for 1.20M and AW = 67. 5 ° .. .

E-VI Vertical Tail Coefficients·at WL 75.0 for 1.20M and AW = 67.5°. ..

E-VII Load Coefficients Data Card Listjng - M=1.20 F-I Wing Coefficients at XRS 3.54. for 0, 9SM and AW = 67. S° . . 67 F-II Horizontal Tail Coefficients at BP 10.75 for 0.95M and AW = 67.5° " . 68 F-III Vertical Tail Coefficients at WL 136.56 for 0.95M and ,,\W = 67. 5 ° . . ...

F-IV Forward Fuselage Coefficients at FS 528.5 for 0.95M and AW = 67. 5°. .. 70 F-V Aft Fuselage Coefficients at FS 1337.5 for 0.95M and AW = 67. S° .. 71 F-VI Vertical Tail Coefficients at WL 75.0 for O;95M and ,,\W = 67.5° . . 72 F-VII Load Coefficients Data Card Listing - M = 0.95 73 Rev Oct 1981 vi LIST OF TABLES (Concluded) Table Page G-I Wing Coefficients at X 354 for O. 70M and AW = 67.5° . 76 RS G-II Horizontal Tail CoefficIents at BP 10.75 for 0.70M and AW = 67.5°. . . . . . . . . . . . .. . ....

G-III Vertical Tail Coefficients at WL 136.56 for O. 70M and AW=67.5°......... . .....

G-,IV Forward Fuselage Coefficients at FS 528.5 for 0.70H and AW = 67.5°. . . . . . . . ......... . 79 G-V Aft Fuselage Coefficients at FS 1337.5 for 0.70M and AW = 67.5° ................... . 80 G-·VI Vertical Tail Coefficients at WL 75.0 for 0,70M and G-·VII Load Coefficients Data Card Listing . . . . . . . .

H-·I Wing Coefficients at X 354 for 1.60M and I\.W = 67.5° 85 RS H-·IIH Horizontal Tail Coefficients at BP 10.75 for 1.60M and H- III Vertical Tail Coefficients at WL 136.56 for 1.60M and H-IV Forward Fuselage Coefficients at FS 528.5 for 1.60M and AW = 67.5°. . . . . . . . ......... . 88 H-V Aft Fuselage Coefficients at FS 1337.5 for 1.60M and H-VI Vertical Tail Coefficients at WL 75.0 for 1.60M and AW = 67.5°, .. " ........ . 90 J-I Wing Coefficients at X 354 for 2.20M andAW = 67.5° 94 RS J-II Horizontal Tail Coefficients at BP 10.75 for 2.20M and J-III Vertical Tail Coefficients at WL 136.56 for 2.20M and A =67.5°......... . ... 96 W J-IV Forward Fuselage Coefficients at FS 528.5 for 2.20M and AW = 67.5°. . . . . . . . ......... . 97 J-V Aft Fuselage Coefficients at FS 1337.5 for 2. 20M and AW = 67. 5 0. • • • • • • ; • • • •• •••••• 98 J-VI Vertical Tail Coefficients at WL 75.0 for 2.20M and AW = 67.5° ........... . . 99 J-VII Load Coefficients Data Card Listing . 100 Added Oct 1981 vii This Page Intentionally Left Blank AIRLOADS RESEARCH STUDY AIRLOAD COEFFICIENTS DERIVED FROM WIND TUNNEL DATA By M. Bartlett, T.F. Feltz, A.D. Olsen Jr., D.B. Smith, and P.F. Wildermuth External Structural Loads SUMMARY This report describes the development of the B-1 aircraft rigid wind tunnel data for use in subsequent tasks of the Airloads Research Study CARS).

The basic intent of the overall ARS program is to utilize flight data acquired during B-1 aircraft test flights, perform analyses of these data beyond the scope of Air Force requirements, and prepare research reports that will add to the technology base for future transport aircraft. Efforts are scheduled as distinct tasks with separate reports for each task.

During this contract phase, existing programs and data from the Rockwell International external structural loads files data bank were used to generate coefficients of rigid airload shear, bending moment, and torSion at specific component reference stations for each of the aircraft aerodynamic effects.

Typical aerodynamic effects for each component include those due to alpha equals zero, alpha, beta, etc. The coefficient data are'presented in slope intercept form. The data presented in this report are for the B-1 aircraft with the aircraft in various wing sweep configurations at selected mach numbers.

I NfROruCTI ON The B-1 aircraft No. 2 (figure 1) is being utilized to conduct the airloads survey flight test program. This aircraft has undergone extensive ground testing to calibrate the strain gages utilized in the airload survey. A com- prehensive wind tunnel test program has been conducted to obtain basic force data and pressure distribution data for both subsonic: and supersonic speeds.

The aircraft provides a reasonable simulation of a future transport aircraft since it has a speed capability in excess of mach 2.0 and employs a large flex- ible structure (figure 2).

The air10ads data gathered from the flight, ground, and wind tunnel tests can be utilized in tlie evaluation of recently developed NASA computer programs, such as NASTRAN and FLEXSTAB, to enhance the analytical techniques of predicting aeroelastic response of large flexible aircraft.

Rev Oct 1981 AIRCRAFT DESCRIPTION The B-1 aircraft is a prototype, long-range supersonic bomber with the capability of high-speed flight at low altitude. Configuration dimensions and general arrangement are presented in figure D-l. The aircraft utilizes a blended wing-body concept with variable~sweep wings, a single vertical stabilizer with a three-section (upper, intermediate, and lower) rudder, and horizontal stabilizers which operate independently to provide both pitch and roll. control. The variable sweep (15 to 67.5 degrees) wing, equipped with slats, spoilers (which also function as speed brakes), and flaps, provides the aircraft with a highly versatile operating envelope. Canted vanes, mounted on each side of the forward fuselage, are part of the structural mode control system which reduces structural bending oscillations in the vertical and lateral axes.

The aircraft is powered by four YFlOl-GE-lOO dual-rotor augmented turbo- fan engines in the 30,OOO-pound-thrust class. The engines are mounted in twin nacelles below the wing, approximately at the left and right wing pivot points. For supersonic speeds, an air induction control system varies the internal geometry of the nacelle inlet ducts to maintain the required airflow to the engines for all flight ·conditions.

Fuel is carried in integral tanks in the fuselage, wing carry-through, and wing outer panels. The fuel system is pressurized and inerted by nitrogen.

Fuel transfer sequencing is automatic and provides center-of-gravity control.

The aircraft has both in-flight and single-point refueling capabilities.

Fuselage The fuselage (figure D-2) is constructed primarily of aluminum alloy materials arranged in a sernimonocoque skin-frame longeron type of construction.

Titanium is used in the wing carry-through, nacelle, and tail support struc- ture, and for various other structures where high load concentrations exist and on the aft fuselage skins where high temperatures and acoustic levels are prevalent. Dielectric materials such as polyimide quartz and fiber glass are used for radornes and antenna covers.

The fuselage structure is fabricated in six major sections and then mated together prior to attaching wings, empennage, landing gears, ~d nacelles.

The following functional description of each section will provide a better understanding of the overall fuselage and its relationship to most of the air vehicle subsystems.

The crew module assembly provides a sealed enclosure with crewmember provisions and is an ejectable unit for emergency escape. The structure is capable of pressurization for a 2,439 meter (8,000 feet) altitude environment mld incol~orates a clear vision windshield designed to bird-proof require- ments, additional crew windows, an entry door, and an emergency exit hatch for ingress "md egress. The floor structure supports crew seating and ej ection rocket loads. An unpressurized section aft of the crew quarters houses the escape system parachutes and provides support for the stabilizing fins. Two sets of deployable mechanical stabilizing spoilers are hinged in the side panel framework and at the lower forward edge of the module. Structural ties to the forward fuselage are severed by explosive charges for emergency escape.

The forward fuselage section includes the nose radome, a forward avionics compartment, an in-flight refueling receptacle, the nose gear well and support structure, a central avionics compartment, a section of the forward fuel tank, a Doppler radome, an environmental control system equipment bay, and the crew entry stairladder structure and mechanism installation. The section also includes many other items of equipment such as antennas and pressure sensiIlg devices. Left and right structural mode control fin surfaces are mounted on this section. Many large and small access doors are provided due to the high density of equipment installations in this assembly.

The forward intermediate section houses the forward and center weapons bays. Maj or bulkheads located between the two bays and at each end of the bays provide support for the rotary weapons launchers. The aft bulkhead location also forms a part of the wing carry-through section. The forward bulkhead is also a closeout for the forward fuselage section. Large integral fuel tmlks are incorporated into the forward intermediate fuselage structure immediately outboard of the weapons bays. A systems routing tunnel occupies the upper structure area between longerons. Provisions for avionics are incorporated in the side fairing area, consisting of equipment bays, antennas, and radomes.

Provisions for external stores pylons, wing sweep actuation components, and flap and slat drive mechanisms are also incorporated in the forward intermedi- ate fuselage section.

The aft intermediate fuselage consists of the main gear well and the aft weapons bay. It incorporates a flight controls mixer compartment and a fuel tmlk above the main gear well. The gear uplock support structure is located in the mixer compartment. Avionics prOVisions are made in the compartment between the wheel wells and in the structural compartments outboard of the wheels. Bulkheads at the forward and aft end of the weapons bay support the weapons bay rotary weapons launcher. As in the forward intermediate fuselage, fuel is stored outboard of the weapons bay. A double support frame for the aft portion of the nacelle extends outboard to the centerline beam of the nacelle. This support is approximately midway between weapons bay

bulkheads. The upper centerline longeron and the lower outboard longerons are located and constructed so as to provide a high stiffness to weight ratio.

The upper centerline longeron extends forward into the wing carry-through section and aft to the vertical stabilizer front spar.

The aft fuselage is a semimonocoque structure and consists of the aft fuel tank area, the dorsal area, the aft avionic bay, and the tail cone. The tank area is closed in the forward and aft end by bulkheads. The forward bulkhead separates the aft fuel tank from the aft weapons bay. The aft bUlk- head closes the tank and provides mounting support structure for the horizontal tail spindle fitting and the aft avionic bay. The dorsal area is a dry tunnel space which houses flight control cables and hardware and provides for the routing of the electrical conduits.

Wing The wing consists of the wing pivot, outer wing panel, flaps, slats, and spoilers. (See figure D-3.) The wing pivot consists of the pin, bearings, and inboard and outboard lugs with provisions for attachment to the wing carry-through fuselage section and the wing outer panel.

The wing outer panel consists of a structural box with leading edge slats, trailing edge flaps, and spoilers over the flap leading edge. The outer wing is mounted on pivot bearings whose supporting lugs are mechanically attached to the wing covers. Provisions for integral fuel containment are provided in the outboard wing structural box. Access is provided for sealing, inspec- tion, servicing, and replacement of fuel system components. Control surfaces on the wings include flaps, slats, and spoilers. The flaps are aft of the wing rear spar and are mounted on rollers between curved tracks. Flap actuat- ing jack-screws are located in the midbay of the flap panels. Segmented leading edge slats are provided. Each segment is supported on tracks mounted on rollers attached to the fixed leading edge structure. Segmented wing spoilers are aft of the wing rear spar and above the flaps.

Nacelle The nacelle is constructed of aluminum alloy, titanium alloy, and stain- less steel and fiber glass laminates. Structural type is semimonocoque with skins, frames, longerons, and a honeycomb sandwich duct. (See figure D-4.)

Each nacelle is fabricated in two major sections, the forward section and the engine compartment.

The forward section consists of the inlet section, the duct assembly, the ramps, and the center beam. The inlet section consists of the center splitter wedge and the upper and lower leading edges. Portions of the upper and lower leading edges are porous for boundary layer control. The duct assemblies con- sist of engine air intake ducts supported by frames and stringers and covered with an external skin. In the forward area, the duct wa1l is covered with aluminwn-machined skin. The intermediate and aft duct wa1ls are covered with fiber glass honeycomb sandwich. The inboard wall of the duct is made up of a fixed duct and movable ramps which provide for a variable geometry system for air induction control. The center beam consists of four main longerons, inter- connecting shear panels, appropriate frames, and the foward nace1le attach point, and is the primary vertical bending member of the nace1le.

Th~~ engine compartment consists of the principal firewa1l bulkhead, the structure between the two engines, the primary engine attach points, and the aft nace1le attach points. Large hinged engine access doors are provided to complete the engine enclosure. Construction is frame, skin, and longeron.

The nace1le is attached to the air vehicle at four points. The forward attach point is a single fitting on the top of the centerbeam structure which is connected to the wing pivot pin through a ba1l joint. The other three attach points are in the engine section in line with the rear engine support.

'Ibey consist of links, two vertical and one horizontal, which connect the nacelle to the heavy support frame extending from the aft intermediate fuselage.

Horizontal Stabilizer The horizontal stabilizer (figure D-5) consists of left- and right-hand slab pmlels attached to a steel spindle projecting out of the aft fuselage stub structure. Both left- and right-hand panels rotate on bearings and are independently controlled in order for the stabilizer to provide pitch and roll control of the air vehicle. Each panel consists of a main structural box, a leading edge assembly, a trailing edge assembly, a tip assembly, an aero- dynamics chord plane seal at the inboard end, and an air seal around the spindle.

Vertical Stabilizer The vertical stabilizer consists of the main box structure, leading edge assembly, tip assembly, and trailing edge structure. (See figure D-6.) The main box assembly supports the two upper rudder segments. Routing tunnels are provided in enclosed areas of the main box structure for electrical and cool- ing lines required to support avionics and antenna equipment located in the tip and leading edge components. The rudder consists of three segments. The upper two segments are attached to the vertical stabilizer through power hinge fittings and actuated by hydraulic motors in the horizontal stabilizer actuator fairing. The lower rudder segment is supported by conventional hinge fittings and actuated by linear actuators between the rudder and aft fuselage structure.

The vertical stabilizer is attached to the aft fuselage principally through a double shear attachment provided on the horizontal stabilizer spindle fitting. The vertical stabilizer is mechanically attached to the spindle fitting by close-tolerance, high-strength bolts.

Stability and Control Augmentation System The stability and control augmentation system (SCAS) provides desired damping and maneuver control. The SCAS transforms pilot pitch and lateral stick displacements and aircraft motion about the pitch and roll axis into symmetrical and antisymmetrical horizontal stabilizer displacements. Similarly, the yaw SCAS employs lower rudder displacement for aircraft motion about the yaw axis.

STRUCTURAL LOADS DATA DERIVED FROM WIND TUNNEL TESTS The basic B-1 rigid aerodynamic data have been obtained from 7,255 wind tunnel test hours involving 14 wind tunnels and 17 models, including those for force measurement, high lift, pressure loads, rotary derivative, spin tests, etc. Data used for structural load analyses are derived from these tests.

The pressure loads wind tunnel model is shown in figure 3. Pressure measurements were obtained by 346 flush static-pressure taps on the top and bottom of both wings and left side of the fuselage (including the wing hood area), the bottom of the left nacelle, the left and right sides of the right nacelle, the top and bottom of the left horizontal tail, and the left side of the vertical tail, including the rudder. Typical wing unit additional lift distributon (CL = 1. 0 due to a) and the corresponding normalized (cl = 1. 0 due to a) chordwise pressure distributions derived from the pressure model test data are shown in figures 4 and 5. The pressure distribution shown in figure 6 was obtained using the data of figures 4 and 5 and normalizing the distributions to produce a wingload equal to unity.

The foregoing data sources and other pressure data contained in the Rockwell External Structural Loads Group's files were used to establish a basic data bank. The pressure data consist of wind tunnel data in the form of span- wise and chordwise distributions versus angle of attack and/or angle of side- slip. Force coefficient data and force data derived from pressure data that are obtained from two separate sources usually do not match. The force coef- ficients computed by integrating the pressure distributions vary slightly in magnitude and center of pressure from the measured data of the force model.

The data taken from the force model were considered more accurate, and the pressure distributions were adjusted to match the force data.

While complete aircraft data are available, it is more convenient (from an external load point of view) to use force coefficient and load distribution data on the individual aircraft components. Data are available on the five basic components; i.e., wing, horizontal tail, vertical tail, rudder, and fuselage. The rudder data are separated into two parts: one for the upper rudder, and one for the lower rudder. The individual component force data reside on a basic data disk pack and may be accessed directly. The pressure distribution data have been converted to "unit" grid loading. This was done by interpolating a pressure map in-the-sTImll and integrating over each grid area. These data also reside on disk and may be directly accessed .. The type of data stored represents the basic data for various aerodynamic effects.

Force and load distribution infollllation for each component are available for the following aerodynamic effects: (1) Alpha equals zero (a = 0) (2) Angle of attack (a) (3) Angle of sideslip (6) (4) Symmetric and antisymmetric control surface deflections (oH' o'H' aSp, oR) (5) Pitch, roll, and yaw rate effects (Q,P,R) (6) Typical cross-coupling terms such as change in sideslip due to change in angle of attack The mUltiplicity of aerodynamic effects results from the use and the arrangement of the aircraft movable and control surfaces which are shown in figure 7. The B-1 arrangement features the following: (a) Blended wing-body concept with variable-sweep wings (b) All-movable horizontal stabilizers for longitudinal and lateral control and trim (c) Spoilers for additional lateral control and for use as speed brakes (d) Single vertical stabilizer with rudders for directional control and trim (e)" Canted movable vanes on the forebody for structural mode control (vertical and lateral) to produce improved ride quality (f) Wing flaps 'and slats to provide improved lift for takeoff and landing The data, as stored on disk, may be used to generate component loads for a complete aircraft flight condition. Distributed load conditions are obtained by applying the appropriate factors to each basic unit loading distribution and summing to obtain a complete set of net aircraft grid loads.

Existing programs and basic aircraft component data from the data bank are used to generate a coefficient of shear (CVi) , bending moment (CBi)' and torsion (CT') at component reference stations for each aircraft aerodynamic effect (i).l These coefficient data are obtained in slope intercept form for use in equations to obtain net rigid airload coefficients for a given flight condition.

METI-IODS AND RESULTS Coefficients of rigid airload shear, bending moment, and torsion are developed at selected airframe stations for the B-1 configuration with selected wing leading edge sweep positions (AW) and selected mach numbers, using the aerodynamic data and methods described as follows.

The force data and the corresponding unit grid load distributions for each component and aerodynamic effect which reside on the Rockwell data bank disk are utilized. The force coefficient data are based on a fixe.d-wing reference area, SRFF' semi span, bREp/2, and mean aerodynamic chord, eREF, which are used for all vhng sweep positions. This reference geometry corresponds to the wing geometry when the wing leading sweep angle (fIw) is 15 degrees. The summation of the wlit grid loads on each individual component equals unity, and the grid loads are normalized using the pressure distribution over the individual com- ponent and its actual geometry. The normal force coefficient for the individual component, CVj' based on the individual component area, Sj' is determined using the wind tunnel normal force coefficient, CNj' based on the reference area, SREP, as follows: (1)

Cy. = C . ( S~p )

N J J J Values of unit shear, bending moment, and torsion, V ' Bu', and Tu " Uj respectively, at any station along the component were obtamed by integrJtion of the unit grid load distribution. The values of shear, bending moment, and torsion coefficients CVij' CBij and CTi' for each aerodynamic effect (i) on the component ~j) were then determined as follows using the individual component,Sj, b /2, and Cj values: j Let: V .. , B .. , T .. = The values of shear, bending moment, and torsion 1J 1J 1J .- Wind tunnel data reference area, semispan, and MAC (for fw = 15°).

C'v .. ' C'B .. ' C'T .. = Shear, bending moment, and torsion coefficient 1J 1J 1J based on SREF' b /2, and C REF REF (2) v..

= 1J = C V C' (3) v..

qSREF N U ij ij 1J

V (_CNi~j SRE_F)

(4) u.. S.

1J J (5) B .. N ..

C)

_ 1J 1J (6) C' B

(

B .. - b pi2 U ..

RE 1J 1J

SREF (b /2)]

Nij REF = B [C SREFJ C = C' (7) B.. B .. [ S. (b./2) U .. S. (b./2) 1J 1J 1J J J J J (8') C' (9) T ..

1J

SREF :REF) = T (\j :REF )

= C' (10) C T .. T..

( . S. C. U S. C.

ij 1J 1J J J J J NOTE: For wing and horizontal tail, all coefficients (CN··' CV··' C ·' C ·' BiJ TiJ etc) are for one side. 1J 1J Table I presents the individual aerodynamic effects which are applicable to wing, horizontal tail, vertical tail, forward fuselage, and aft fuselage compo- nents. Shear, bending moment, and torsion coefficients have been calculated at one station on each of the preceding listed components.

Appendix A presents a list of symbols used in this report and their definition and units where applicable. Figure B-1, of appendix B, presents the location of the reference points for which load coefficients have been determined.

The resulting coefficients of rigid airload shear, bending moment, and torsion at each component station for each of the applicable aerodynamic effects listed in table I are presented in the appendixes for the various mach numbers along with their applicable areas, semi spans , and mean aerodynamic chords.

Several of the aerodynamic effects must be combined to obtain the net rigid air load coefficients for a given flight condition. The equations in appendix C present the method of combining the aerodynamic effects at each of the stations. Refer to the list of symbols (appendix A) and appendix B for the definition of symbols and dimensional data.

Some anomalies in the equations of appendix C should be explained at this point. Aircraft Sideslip (13) produces an asymmetric load distribution on the wing. To accommodate this, the asynnnetric distribution was separated into an equivalent symmetric and antisymmetric part such that when added together pro- duces the asymmetric load distribution. The force data are aiso treated in this manner. The sideslip also produces a cross-coupling effect and varies with changes in angle of attack. This method of treatment can be seen by examining the equations. Due to the asymmetric type of load distribution ey~ibited by some aerodynamic effects, C ' CB' and,CT are given for both the V left and right sides for the wing and horizontal tail.

The fuselage has both vertical and lateral load distributions, and separate Cv and CB equations are presented. The net airload coefficient equations for the aft fuselage include the airload on the aft fuselage and the air10ads on the horizontal and vertical tail surfaces.

The rigid airload coefficients of tables B-1 through B-VI for mach 0.85 and sweep 67.5 degrees are prepared in punched card form, as will all subse- quent rigid air load coefficients for the various mach number-wing sweep com- binations. The data cards have been formatted to be compatible with the NASA DFRC Cyber 73-28 computer. A listing of the data cards is presented for mach 0.85 (table B-VII), and will be presented in subsequent appendixes for all subsequent mach number-wing sweep combinations.

The total block of data cards contains 12 descriptive title cards. Ninety data cards follow the 12 descriptive title cards. The data cards are in six groups of 15 cards each. Each group represents the rigid airload coefficients for one of the five stations where loads will be measured during the B-1 flight loads survey plus the vertical tail root station. Each data card contains a sequence number which is explained in the descriptive title cards, C ' CB, C , T V and an aerodynamic effect description for each set of CV' CB' CT. The filler cards provide space for up to 15 aerodynamic effects at each of the six reference stations.

The suggested format for reading the rigid aerodynamic coefficient cards into the NASA DFRC Cyber 73-28 computer is as follows: • Descriptive title cards (12 cards) Format (8AlO) • Rigid aerodynamic coefficient data cards (90 cards) Format (110, 3ElO.2, 4AlO) ACKNOWLEDGMENT Mr. R. Celniker is recognized for his important contributions as program manager during the early parts of the study.

REFERENCES 1. Rockwell International Report TFD-72-10l7, "B-1 Rigid Aerodynamic Data for Stabilty and Control, Status at Air Vehicle DVR (As Revised to Present Final Preflight Data)," 25 November 1975.

2. Rockwell International Report TFD-73-960, "B-1 Aerodynamic Force, Moment and Load Distributions," 12 September 1975.

3. Rockwell International Report NA-71-522, "Transonic Wind Tunnel Test of the 0.036 Scale B-1 Pressure Loads Model," (TWT 236), 22 June 1971.

Rev Oct 1981 13 TABLE I.- AERODYNAMIC EFFECTS APPLICABLE TO CO~~ONENT LOADS Horiz Vert Fwd Aft Effect Wing tail tail fus fus 0- X X X X = 0 0- X X X X .

X X 0- X X S X (sym horiz tail defl) X °H horiz 0' H (anti syrn tail defl) X X X 0SP (spoiler defl) X X X 0SP c/o (horiz tail carryover) X (upper rudder defl) X °RU (lower rudder defl) X X ORL P (damping in roll) X X X X X (damping in pitch) X Q X R (damping in yaw) X

8a. = o A/S (wing) X

=

o Syrn (wing) X

8a.

BaA/S (wing) X SO-Syrn (wing) X = 0 (vert tail) X 8a.

(vert tail) X Sa.

= 0 X 8a. c/o (aft fus carryover) fus carryover) X 8a. c/o (aft X = Applicable aerodynarnlc effect Figure 1. - B-1 aircraft.

Figure 2. - Structural breakdown.

Figure 3.- Wind tunnel pressure loads model.

.......

1.6 I I 1.0 =

l(cC! C ~d~

V""

o L avg

~

..-/ 1.2

.......-

-

-

~

c 1. c

"

c . c .8 ......

L avg

"

~

~

.4

\

o o • 1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1] Figure 4.- Unit additional span load distribution on wing - centerbody1\W = 67.5°, M = 0.85.

1 .O/CPL c-" CPu) dX/C \ ~ 1.0 at each l1-station

1 4~------~------~------~------~-- 0

'Y/=O.O

.. ............::.---

o • 1 .2 .3 .4 .7 1.

.5 .6 .8 .9 X/C o Figure 5.- Unit additional chordwise load distribution on wing-centerbody llW = 67.5°, M = 0.8S.

WING PRESSURE PLOT 67.5 DEGREES SWEEP n 128 Ma .85 SW 67.5 ADDITIONAL SYM RIGID DIST 558 WING Z SCALE ZMAX ~.59XIO-05 ZMIN -8.77XIO- I INCH 2 2.02XIO- Figure 6.- Normalized pressure distribution.

RH hori zonta 1 stab i 1 i ze r RH spo i 1 ers ~~~~l;:---- Mid rudde r Lower

er

\.rUdd

LH ho r i zon ta 1 stabi 1 i zer Wi ng sweep LH spoi lers

/

SMCS vanes LH flaps (6) (canted 30 de ) 9.

LH slats (7) control surfaces.

Figure 7. - Flight

APPENDIX A SYMBOLS wing sweep angle degrees q c:lynamic pressure Ib/ft Z component area, reference area ft Sj, SREP component semispan, reference semispan in.

bj/Z, b /2 REp mean aerodynamic chord, reference mean in.

aerodynamic chord (MAC) normal force coefficient based on C v.

J component area normal force coefficient based on reference area unit shear, bending moment, and lb, in. -lb, torsion on component j in. -lb B· T· Shear, bending moment, and lb, in. -lb, v· 1. 1. 1· J , J , J torsion on component j due to in. -lb aerodynamic effect i C C C coefficient of shear, bending moment, B .. T ..

V ..

1J , 1J , 1J and torsion on component j based on aerodynamic effect i using component area, s emi span , and MAC C' C' C' coefficient of shear, bending moment, v .. B .. T ..

1J, 1J , 1J and torsion. on component j based on aerodynamic effect i using reference area, s emi span , and MAC angle of attack, + nose up degrees .

degrees/sec r~ rate of sink, + nose up angle of yaw, + nose left degrees degrees/sec p rolling velocity, + left wing up degrees/sec Q pitching velocity, + nose up degrees/sec R yawing velocity, + nose right horizontal tail deflection, + leading degrees edge up differential horizontal tail degrees «5' H deflection, +o'H produces a plus rolling moment .cleft wing up) spoiler deflection + when right degrees spoilers deflected effect of spoiler deflection on degrees horizontal tail + when right spoilers deflected deflection of upper segment of degrees rudder, + trailing edge left deflection of lower segment of degrees rudder, + trailing edge left o A/S effect of ~ on wing, anti-symmetric degrees contribution, + nose left A/S effect of ~ and on wing, anti- degrees-degrees symmetric contribution, + nose left, + nose up *~ry = 0 Sym effect of ~ on wing, symmetric degrees contribution, + nose left Sym effect of ~ and on wing, degrees-degrees symmetric contribution, + nose left, + nose up ~(X = 0 c/o effect of ~ on aft fuselage, degrees + nose left effect of ~ and on aft fuselage, degrees-degrees c/o + nose left, + nose up *The explanation as to why the aerodynamic effects are divided into antisymmetric and symmetric parts is found in the methods and results paragraphs of this report.

A - airplane c£ - section lift coefficient c - section chord - C average surface chord avg - rj fraction of semispan - pressure C coefficient p - L lower surface - upper U surface

APPENDIX B RIGID AIRLOAD COEFFICIENTS AT LOAD REFERENCE POINTS OM = 0.85) The location of the load reference points where the rigid air load coefficients have been determined are presented in figure B-1. All of these stations, with the exception of the vertical tail root (VTR) , are the stations where loads will be measured during the B-1 flight loads survey program.

The rigid airload coefficients were determined for each of the applicable aerodynamic effects listed in table I and are presented in tables B-1 through B-VI. These coefficients were determined using equations 4, 7, and 10. The component applicable reference areas, semispans, and mean aerodynamic chords are listed in the coefficient tables.

TABLE B-I. - WING COEFFICI~'TS AT XRS 354 FOR 0.85M AND l\'{ = 67.5°

1946.0 ft

=

Sw 820.08 in.

= b\!2 _.

C 184.05 in.

w

Coefficients are applicable to either left or right wing • .

-

C Crr (~Z BX (torsion) Effect (shear) (moment) --- , I a 0 .024108 = .006613 .000163 .009900 a .002424 .000861 • .005085 a .001245 0000442 -.000357 ~('P -.000089 0000019 ,) - .002572 P -.000748 -.000147 .043629 Q .012035 -.000:575 So. = o A/S -.001022 -.000189 .000088 S().

-.000234 -.000056 -.000018

A!S

.000058 -.000008 .000041 SCt o Sym

=

Sym -.000099 -.000025 -.000005 Sa .

+ Up and perpendicular to the wing reference plane.

c;,Z'

+ Tip up and about an axis perpendicular to the wing load reference line (0.36c line)*.

Leading edge up and about the wing load reference line (0.36c line)*.

*

The wing load reference line passes through the pivot (at XRS 139.51S, YRS -49.845) and the load reference point (at XRS 354, Y -38.248).

RS Rev Oct 1981 tABLE B-II. ~ HORIZONTr\L TAIL COEFFICIENTS AT BP 10.75

EOR 0 i 85M AND 1\1[ = 67. 50

= 238.77 ft2

%r

b /2 259.03 in.

= HT 149.38 in.

=

S-rr

Coefficients are applicable to either left or right horizontal tail.

C

BX CrY Cvz

Effect (shear) (moment) (torsion)

= a

a -.167555 -.066791 0029355 a .042901 .018742 -.013642 ~H .077595 0033898 -.024675 .

a .248221 0108438 -.078933 -.008687 S -.002984 .001240 ~'H .027446 .012417 -.008247 -.000717 (Sym) ~SP -.000352 .000269 (a/s) ~SP .000268 .000179 -.000164 p -.002504 -.001927 .002056 Q .2H383 -.204014 .495028 + Up and perpendicular to the airplane water plane.

CyZ'

+ Tip up and about an axis parallel to the longitudinal axis.

Cyy, + Leading edge up and about an a..'<:is perpendicular to the

plane of synnnetry.

TABLE B- II I. - VERTICAL TAIL COEFFICI~'TS AT WL 136.56

FOR O.85M AND J\v = 67.5°

247.4 ft

=

~ 2 = 206.76 in.

hvr/

188095 in.

=

CVr

Coefficients are applicable on the upper vertical tail 0JVT).

Cyy CBX

GrZ

Effect (shear) (moment) (torsion)

-

-.030790 -.009622 .000175

Sa = 0

-.000265 .000048 Sa -.000847 ., -.003621 .000784- -.013635 ~H l~p -.000226 -.000067 -.000002 ~)RU .014456 .004492 -.004046 ~)RL () p - .00:3048 -.001124 .000367 R .024764 .007752 -.002350

-

C , + T6 the right ruld normal to the plane of s~try.

Yi + Tip to the right and about an axis parallel to the

S3x'

longitudinal axis.

L- + Leading edge right and about an axis perpendicular -TZ' to the water plane.

TABLE B-IV.- FORWARD FUSELAGE COEFFICIENTS AT FS 528.5 FOR 0.85M AND AW = 67.5° SFF = SREF = 1946.0 ft bFF/2 = bREF/2 = 820.08 in.

184.05 in.

CFF = CREF Coefficients are based on the air10ads on the fuselage forward of FS 528.5.

Cvy CTX Cvz CBY CBZ (moment) (torsion) ,Effect (shear) (moment) (shear) - - - a = -0.000068 0 0.006532 - - - 0.000553 a 0.001730 p - - 0.000168 0.000044 0.000026 - -0.000356 - -0.002047 -0.000684 S CVZ' + Up and normal to the water plane.

CBY, + Nose up and about an axis perpendicular to the plane of symmetry.

Cyy, + To the right and normal to the plane of symmetry.

CBZ, + Nose right and about an axis perpendicular to the water plane.

CTX, + Left wing up and about an axis parallel to the longitudinal axis.

TABLE B-V.- AFT FUSELAGE COEFFICIENTS AT FS 1337.5 FOR 0.8SM AND AW = 67.5° SAF _. SREF = 1946.0 ft bAF/2 =: bREF/2 = 820.08 in.

184.05 in.

Coefficients are based on the airloads on the fuselage aft of FS 1337.5 and do not include the airloads on the empennage. (Refer to appendix C for equations which include the airloads on the empennage.)

CVZ CBY C CBZ CTX VY Effect (shear) (n loment) (sh ear) (moment) (torsion) ..

( - - a = 0 ), 003220 0.009376 .- - a - -0.000181 - 0.000052 - So. = ··0.0 -0.000241 -0.000290 0 c/o 01670 - So. c/o ··0.0 00046 -0.000007 -0.000008 O'H - 0.000239 0.000158 0.0 00910 - 0.0 00153 0.000021 0.000027 °RL p - 0.0 00541 0.000133 0.000094 - 00334 -·0.000068 -0.000058 B -0.0 -.

CVZ' + Up and normal to the water plane.

CBY, + Aft end up and about an axis perpendicular to the plane of symmetry.

Cvy, + To the right and normal to the plane of symmetry.

CBZ, + Aft end right arId about an axis perpendicular to the water plane.

CTX, .t- Left wing up aIld about an axis parallel to the longitudinal axis.

J'ABLE B-VI. - VERTICAL TAIL COEFFICIDrrS AT WL 75.0 FOR 0.85M AND "~Ij = 67.5° ft

= 247.4

~ 2 206.76 = in.

bvr/ 188.95 in.

=

CVr

Coefficients are at the vertical tail root (VIR) and are for use in the equations in AppendLx C for the determination of the net air10ads coefficients for the aft fuselage point at FS 1337.5.

C C vy BX

Crz

Effect (shear) (moment) (torslOn) -.020655 .007677 SCI. 0 -.046314

=

-.000568 .000211 -.001274 SCI.

~'H -.007032 .003313 - .007260 -.000142 .000035 ~SP -.000310 .008801 oRIJ .014473 -.007605 .006167 .000570 -.000897 0ru..

-.002042 .001309 p -.002933 .016997 -.007621 R .037307 + To the right and normal to the plane at symmetry.

lw'

+ Tip to the right and about an axis parallel to

S3x'

the plane of synnnetry.

Cyz' + Leading edge right and about an axis perpendicular

to the water plane.

TABLE B-VII. - LOAD COEFFICIENTS DATA CARD LISTING - Concluded TABLE B-VII.

wing LW, RW - left and right at X 354.0: ~ RS < n in. 354.00 in.

BP ± 240.672 XRS ct or Y -38.248 in.

FS 1161.365 in.

f-' RS t.D 4.370 in.

WL 9.075 00 ZRS f-' horizontal ta i 1 : LHT, RHT - left and right BP ±10.75 in.

FS 1582.0 in.

WL 126.0 in.

- UVT upper ta i I: verticaJ WL 136.56 in.

FS 1582.0 in.

BP 0.0 in.

- vertical tai I VTR root: in.

WL 75.0 NOTE: REFER TO TABLES B-1 THROUGH in.

FS 1535.56 B-VI FOR DEFINITION OF BP 0.0 in.

REFERENCE AXES.

- FF forward fuselage: FS 528.5 in.

WL 32.0 in.

BP 0.0 in.

- AF aft fuselage: WL FS in.

i337.5 WL 34.0 in.

BP 0.0 in.

Figure B-l. - Component rigid airload coefficient reference points, AW = 67.5 degrees.

APPENDIX C NET RIGID AIRLOAD COEFFICIENTS FOR FLIGHT CONDITIONS When it is desired to obtain the component net rigid airload coefficients for a flight condition, several of the individual aerodynamic effects must be combined. Equations which are applicable to each load reference point are pre- sented in this appendix. These equations' utilize the coefficient data and geometry data of tables B-1 through B-VI.

Particular care should be exercised in the application of the dimensional units for certain terms in the equations. Refer to the list of symbols (appen- dix A) and appendix B for the definition of symbols, their units, and dimensions.

For example: b/Z and C are in inches.

Area, S, is in square feet.

Velocity, V, is in feet/second.

a, 6, 0H' o'H' eSP' and 0RL are in degrees.

P, Q, R, and a are in degrees/second.

Equations for Wing Station ~ 354 Equations Cl and CZ are generalized equations for the net rigid airload coefficients C ' CBX, and CTy at the left and right wing stations, respectively.

VZ The individual values of CVZ, CBX, and CTY for each aerodynamic effect, appli: cable to theA = 67.5° and M = 0.85 condition, are presented in table B-1.

W w

r. _ r. + r. a + C (a C ) - r. 0

vBTLW VBT =0 VBT VBT· 2V VBTo SP LW a a a SP (Cl) C + (C~ + VBTSa=O -VBTSa

+ [~Tsa=o + ~T )a]s

SaA/S A/S SYM SYM ~ C~)

(

~TRW = ~Ta=o + cBVTaa + ~Ta Zv- + ~To

SP C (C2) VBT

+ [~Tsa=o -

Sa=O SYM A/S Equations for Horizonta~ Tail Station BP 10.75 Equations C3 and C4 are generalized equations for the net rigid airload coefficients CVZ, CBX, and CTY at the left and right stations, respectively.

The individual values of CVZ' CBX, and CTY for each aerodynamic effect, appli- cable to the fl.W = 67.5° and 0.85 M condition, are presented in table B·-II.

(C3) C_ - C + C a, + C .~ + C H 'VETRI-IT VBTa,=a 'VET BVT u BVT· a, . 0H a, 8' H - C- S + C- 6 - L_ 8 T T SP -VB 'VE R - VBT.1' SP L S 8 <J Sp SP c 0 / {C4) Equations for Vertical Tail Station WL 136.56 Equation CS is a generalized equation for the net rigid airload coeffi- cients Cyy, CBX, and CTZ at WL 136.56. The individual values of Cyy, CBX, and CTZ for each aerodynamic effect, applicable to tha AW = 67.5° and a.8SM condi- tion, are presented in table B-111.

CVBT1Tf = [CVBTBU=O + CVBTBUa]B + CVBT ' o'H

O H VBTo +C (osp + aSp ) + ~T °RU SP R L 0RU (CS) Equations for Forward Fuselage Station FS 528.5 Equltions C6 through CIO are for the net rigid airload coefficients CVZ, CBY, Cvy, CBZ, and CTX at FS 528.5. The individual values of CVZ, CBY, cvy, CBZ, and CTX for each aerodynamic effect, applicable to the AW = 67.5° and 0.85M condition, are presented in table B-IV.

(C6) (C7) pF (Pb ) = (C8)

~BB + ~P --zv-

~PP

pF (Pb ) = C (C9)

C S S + C --:zv

BZ BZ BZp pF pp ( Pb ) C + C a = (ClO)

+ CTXSS + CT~ -zv

TX TX CT~p a=O a Equations for Aft Puselage Station PS 1337.5 Equations Cll through CIS are for the net rigid airload coefficients CVZ, CBY, W'{, CBZ, and CTX at PS 1:537.5. The net coefficients include the airload on the fuselage, aft of FS 1337.5, and the airloads on the horizontal and verti- cal tail surfaces. The empennage airloads are included by using the horizontal and vertical tail root loads and their transfer distances to PS 1337.5. In the equations that follow, the numbers subscripting th~ brackets, { }, denote the table number or the equation number from which the coefficients within the brackets are obtained; i.e., { }BV denotes coefficients are from table B-V { }C3,C4 denotes coefficients obtained using equations C3 and C4.

(Cll)

C = jC + C al

By By By

AF I a=O a B-V

+ [(~Zurr + ~ZWIT) (::b;Z)

(C12)

- (SYurr + SY~IT) (s;rb;zj} C3,C4

(C13)

C = I (C + C a)s + C , 6'H

BZ BZ BZ BZ AF Sa=Oc/o Sa 6 H c/o

+ C 6 RL + C (P~~)+ C s}

BZ BZ BZ 6 P S B-V RL

+ CyyVTR (::~)z) - C (-S;b;z1} CS,B-VI* (C14)

TZVTR *Use equation C5 with coefficients from table B-VI.

C =! C + C a + (C

TX TX", __ O TX TX AF "" a {3Clt=O / C 0 + C

TX '

o

H

I'-, Y rurr Sl-rrj

CvZ ) S C ( rurr AF AF (C15) Moment Transfer Arms I'-,XHT ,I'-,YHT, I'-,XVTR, and I'-,ZVTR The aft fuselage coefficients are determined at a point at FS 1337.5, WL 34, and BP 0.0. The left and right horizontal tail root coefficients are determined at points at FS 1582, WL 136.56, and BP ±10.75. The vertical tail root coefficients are determined at a point at FS 1535.56, WL 75, and BP 0.0.

The corresponding moment transfer arms are then as follows: I'-,~ = 1582.0 - 1337.5 = 244.50 in.

I'-,Y = -I'-,Y = 10.75 in.

rurr LHT I'-,~R = 1535.56 - 1337.5 = 198.06 in.

I'-,ZVTR = 75.0 - 34.0 = 41.0 in.

:~Use equation C5 with coefficients from table B-VI.

APPENDIX D

FIGURES USING ENGINEERING UNITS

--269.03 -- .....

62.87 /,' HORllC\TAl TA~L H: ... GE U~E~ ~/' Yf1582.J ITEM AR"A SQ. FT. 11.5 t.SPc.::T RAT ,0 9.6 3:~ 3.95 1.2 25 'APER h;..T:J .35 .30 .30 .20 .10 ;:'>-Zl~ TH CKNESS K:":-.O ~~:-: ..... :;c5 ::''::.-:. .O~ .0:' ;:= <0'."] ":'0 'l';~ I--Al-R-FO-:-L-S-E-:-~-:':)-~-' --I \;,t.. 69 - i 9:' 2 : 5-;:. ~ - t~Jl::J ""Z- ~':l"X:: Zr; 2-:'_,) ro - P 45' AT .25: __ ~L::: ~"'.:. 0) o· -30.0' o· S=FL ± 22;.C/ 168.9:>4 118.373 84.825 X 3/,<i 2327 F Y S4! 8'..0::.13 F ""'\'I!;"';(, P.VpT XF-'4S0- Y 969.0 F AJRCRf...F- T Er,~Pl Y NEi3HT ....... LB:: --- ~~~!;:S~M8_7 DESIGN USEf'JL LOA) La ~ ---- -' -_ ~~~/~"'8-7 LB -;::- 3SC,OOO· --- DESIGN GROSS 'i![IGH1·;,\XI LB = 391,000 MAXI~'UM GROSS WE!C'" T

.I

/

_.25 CS~( Y,232.IO @ / / ::C-". r 8.579 ref)" f'-:'< /

40.1796--.., ? / "" I

• / ~ / 1_- 42.894 ~32171 -Y 22.0 / F @ yro.o ~----1905.36 (152.78') -- t------------------------L---.:.."LI--164017 0366aJ w,r;G FOR.·,ARO--------J L--------------~ --

1-------------- i -939.0 (7823) WING AFT---- I

'I~ ;1 I 1

~ 100 ~ leo ",! I t I ~ SCALE - IN:>-'oS ZFOO-~ --

~~:::=k:. ====._·~.~~~~~~~IC~~. -~~W===· .-.=:::::::- Z,3====:.5 ======:J=r

1,J - I 107.75

------------J...iOs;!l.C'.LJ1 111L[-_L. __ ~:.7________ --_I

• 1 174_lC~· '\': -'

This Page Intentionally Left Blank

2 3 5 6 7

12 14 16 NOTE Shaded area indicates fuel tanks.

AFT INTERMEDIATE FORWARD STORES STORES STORES BAY BAY BAY

47 43 41 37 36

39 38

Figure D-l.- Continued.

35. LOWER RUDDER 1. PITOT·STATIC BOOM (WITH AOA 47. FORWARD AVIONICS COMPARTMENT AND SIDESLIP VANES) 36. AFT RADOME 48. EJECTABLE CREW MODULE 2. FORWARD RADOME 37. AFT AVIONICS COMPARTMENT 49. FORWARD CREW STATIONS 3. AERIAL REFUEL RECEPTACLE 38. LN2 DEWAR 50. CREW SEAT (4) t 51. ESCAPE HATCH (SEVERABLE) 4. PITOT·ST A TIC PROBE * 39. ENGINE NACELLE * 5. TOTAL TEMPERATURE PROBE * 40. MAIN LANDING GEAR * 52. AFT CREW STATIONS 6. STRUCTURAL MODE CONTROL SYSTEM 41. AERIAL REFUELIWING INSPECTION LIGHT. 53. CONTROLS FOR ENTRY LADDER. APU, AND VANE * 42. ENTRY DOOR MAIN GEAR DOORS 7. ANGLE·OF·ATTACK VANE * 43. ENTRY LADDER 54. SURVIVAL~EQUipMENT 44. NOSE LANDING GEAR· 8. CR'EW ENTRY WAY 55. SIDE WINDOW (SEVERABLE) • 45. LANDING/TAXI LIGHT 9. FORWARD FUSELAGE FUEL TANK 46. LANDING LIGHTS (2) (TANK NO.1) 10. FORWARD INTERMEDIATE FUSELAGE FUEL TANK (TANK NO.2) 11. MAIN FUEL TANKS 12. MAIN WHEEL WELL EQUIPMENT (INTERMEDIATE AVIONICS) COMPARTMENT 13. AFT INTERMEDIATE FUSELAGE FUEL TANK (TANK NO.3) 14. AFT FUSELAGE FUEL TANK (TANK·i')IO. 4) 15. HORIZONTAL STABILIZER ACTUATOR * 16. HORIZONTAL STABILIZER 17. FLIGHT CONTROLS MIXER BAY 18. ENGINES *

19. FLAPS (6) *

20. SPOILERS/SPEED BRAKES (4) * 21. FUEL JETTISON OUTLET * 22. POSITION LIGHT * 23. SLATS (7) * 24. WING FUEL TANK * 25. APU*

t----42

26. HYDRAULIC RESERVOIRS* 27. INLET RAMP MECHANISM * 28. WING PIVOT 29. SUPPLEMENTAL POSITION AND ANTICOLLISION LIGHT * 30. WING GLOVE AVIONICS COMPARTMENT * 31. CENTRAL AVIONICS COMPARTMENT 32. VERTICAL STABILIZER 33. TAIL/ANTICOLLISION LIGHT 34. UPPER AND INTERMEDIATE RUDDERS ~--43

* Both Sides (L and R)

t Right aft seat temporarily B 1 1 048 removed Figure D-l.- Concluded.

- .

2.

I 20

9 6

8 !

" ,'; c ..

I I I

i ! l

. '.

. "" .. , .. , .' .

," ) .... n

, . ' ...•• :' ...... I'.' , '; .,: :1 : ''''

~4 .!' I' •• :,: "'. , •..• , .•.. " •.. ,. • ,". 'I"! ,: i?

. ." ...! .. ' .,

.: ~ I,' 'i'

_, :: I. " :""

, j , .' j ·'Ll. ...•. .." .. "

I

. . vi..: •• ..' .." .'" ' ", ." =, . • r:)/

--;. 'ji" V

i

, i

! , , '---+- . -+-----. ----j-- ---- -' t-f;~z:

VIEW lOOKING DOWN !

II . I , . '. I' . '- 1'-.. ',)!; •. ; 'i '

""~.) '.1 I' '" fj , , !1

:: I I: ;' i,lJ··lil ,," " .•...

, ......

:1 I: L '+;!'+I".;.~.i_.

: : ..•. , ' '., " 1" ••.•. ," , I: ,,' '! ":i:' :, :i: ,,: : I : ;' .. : ,.

" ;1

I'

1 !,' ' .• '. ' ",: ,;: ,'I' "

': "

.. 'I

i ' .. 'ei.o i' '.'. _U . ·i. i

.' , ,

, .. , .. I

: tt' ·.C' L:=. ·L "'-~', '

r-------------+-----~---I~_-~~=M~==~~-L-------------+_------~-F~~~rwa~«;~K=~-·~----_t--~~~1~;~~--------t_·---~'~~~fl~J/~=

~b

--+-- ----j---. -t--- ..

---1. __ , , .

i j ! ~" i~

" '.'

Ii •. ,I ; " , ,.

. .

, 'i :': I ;

,i i'-"

,I ,', { . ..

" .' ..

• ,i • - , • "

L

2.

16 I 13 4

i

~lgure D-t.- Fuselage s~ru~tuTe dlagr&~ (B-1J (-SSE),

This Page Intentionally Left Blank

SPO/U/l /JeT ,()C.4TICW.$ AIf$ X$P (i) dOS.37" 403.4l7 .!FVI/.l/l ./ 43Z.07Z Q) 4"0. III y,., 0

L

y, ~49.S3' 49t.07'! 440. III

Jl

1.

-

~ 417. 0'5 475.08Z

\

StJ3.7~Z.

~ .501. BOB 513.7('1!

~ SI/.8~

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(Ir,s IS 7.)

I

(t) 547.US .545: 2139

i

~ 513.7ZtJ SPCIL£R*~ 57J.7t.'

® 58~7ll ~lJl.763

I

tolS-804- 613.8fJ4 F ~it.OOO ---I- ~ MZ.04C. .5POIL.EI? - 4

".w, C65 y", 3l00UJ

(.35.SSSI %) ~S2.050 6SO.0M

I

I

o

(5

o

N

~

Figure D-3. (-SSB).

This Page Intentionally Left Blank

I

I

-t-

r

r 383.69

Mc<'"~~r L",~~ .J;"';''''''c,- (,..-" ;: ..... w, v,..

,~ ns

v )

CP(l~

T

!

I !

i I

I I

I

__ ",~ d_ ~ :;r"".A1VR4, ,,;,~·.·.p·""'-vr. I/-

.r, -" __ ~., C~r-H'~:"", .i;"rr ..... r:~~ .. 1 [9320085

Figure D-4.- Structural arrangement

- nacelle external compression inlet rRDl ~;:")

,-. ]. '"1-'--'

This Page Intentionally Left Blank

I

+

G·G

VlEIt' L.E YIRT.

------ --

!~

I .. A.;,rraF IkJIf'IZ ~'-'.

LZICXJ()()I R4Nftr 1/$$'( L.!OODOZ

IICW DaD

1ftI!M.I:1.:I

- STRL arrangement

LE sweep)

This Page Intentionally Left Blank

"'" 0 :-,. ,co' ,,,. "-c' ... ,0(;..''' """',A ,,'~ .,,,, " •• (i) =. ,., 'L ,,.,,, 'C~·""'''''i'' @ ",!", Teo,,, •• :"0'''''_ '~O'4, /,\ o ",. -'~ 'c ,~".

o ~J~ .:;" :,' t~,',. 'OC'<)~ ".,.,t(o CO" );lo"o.w;"1

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"''''''. F-F -.u.J., l-i 7~( ~o, ...

V""'"

, '''~ -d V'IE'I'I" OF sp"r;tl\.E F,:T,NG ARO:A 10("'-1.

;.EA"",. (OO! ,~'" <'" < q'fI~' , ~~ , I' I ' "'-'--'- "","" _ ,N( '.,

,., ."

oo.u. cCU'''''' ..c"(~

.... "

",,"«,,>0" i ,

------ ---FRL

Figure D-6. Vertical stab'i" l lzer structural arrangement.

This Page Intentionally Left Blank

APPENDIX E RIGID AIRLOAD COEFFICIENTS AT LOAD REFERENCE POINTS ~ = 1.20) The location of the load reference points where the rigid airload coeffi- cients have been determined are presented in figure B-1. All of these stations, with the exception of the vertical tail root (VfR), are the stations where loads will be measured during the B-1 flight loads survey program.

The rigid air load coefficients were determined for each of the applicable aerodyrlarnic effects listed in table I and are presented in tables E-I through E-VI. These coeffici~nts were determined using equations 4, 7, and 10. The component applicable reference areas, semispans, and mean aerodynamic chords are listed in the coefficient tables.

Notes in appendix C which apply to tables B-1 through B-VI for m = 0.85 data apply correspondingly to tables E- I through E,.. VI for m = 1. 20 data.

TABLE E-I. - WING COEFFICIENTS AT XRS 354 FOR 1.20B AND j\~ = 67.5 ft2 1946.0 = Sw 820.08 in.

= bw'2 184.05 in.

= ~ Coefficients are applicable to either left or right wing.

C Crr Cyz BX Effect (shear) (moment) (torsion) I -.003133 0 .029213 .008831 a = .003229 .000584 .012051 a • 0 0 a .000034 -.000290 -.000057 ~SP -.000955 -.000223 -.003232 P .012062 -.001284 .042987 Q -.000167 .000149 -.000759 Sa o NS = -.000058 -.000025 -.000226 Sa NS .000119 -.000174 .000311

= o Sym

Sa -.000011 .000005 .000008 Sym Sa + Up and perpendicular to the wing reference plane.

CVz'

+ Tip up and about an axis perpendicular to the wing load reference line (0.36c 1ine)*.

Leading edge up and about the wing load reference line (O.36c 1ine)*.

*

The wing load reference line passes through the pivot (at XRS 139.515, YRS -49.845) and the load reference point (at XRS 354, Y -38.248) RS Rev Oct 1981 'tABLE E-II. - HORIZONTAL TAIL COEFFICIENTS AT BP 10.75

EOR 1.20M AND Aw = 67.5°

~ 238.77 ft b /2 = 259.03 in.

HT ~ = 149.38 in.

Coefficients are applicable to either left or rifitt horizontal tail.

C

BX CrY

Cvz

Effect (shear) (moment) (torsion).

a 0 = -.136182 ":'.049482 .051116 a .042237 .018920 -.019427 ~H .060513 .027107 -0027833 • a .233930 .104789 -.107545 (3 ~. 017738 -.006993 .004648 ~'H .049585 .022735 -.022025 (Syrn) -.000317 -.000'750 ~SP .000369 (a/s) .000145 ~SP .000277 -.000168 p -.002088 -.002479 0002313 Q .552160 ~ 2521971 ··.297065

Cyz, + Up and perpendicular to the airplane water !plane e

C ' + Tip up and about an axis parallel to the lImgitudina1 BX axis.

sY, + Leading edge up and about an axis perpendirular to the

plane of symmetry.

TABLE E-III. - VERTICAL TAIL COEFFICIENTS AT WL 136.56

FOR 1.20M AND ~ = 67.5°

247.4 ft =

\rr

2 206.76 in.

=

\rr/

188 0 95 in.

= ~ Coefficients are applicable en the upper vertical tail 0JVT).

CBX Cyy

Crz

Effect (shear) i (moment) (tersien)

Sa = 0 -.010927

-.034484 .003146 Sa -.001115 -.000354 .000102 ,

a

-.003032 -.000800 .000464 H ~p -.000270 -.000089 .000023 ~RU .009663 .002854 - .003830 ~RL 0 0 0 -.001461 .000756 P -.003995 -.005596 R .032297 .010244 + To. the right and nermal to. the plane of syrmnetry.

Cyy, + Tip to. the right and abeut an axis parallel to. the

Ci3x'

lengitudinal axis.

Crz' + Leading edge right and abeut an axis perpendicular

to. the water plane.

TABLE E- IV. - :FORWARD RJSELAGE COEFFICUNrS

AT FS 528.5 FOR 1.20M AND 1\; = 67.5°

_.

1946.0 ft SFF _.

2 820.08 in.

})FF/ _.

184.05 in.

~F Coefficients are based on the airloads on the fuselage fOTIvard of FS 528.5.

C C Cyy CBY Crx VZ BZ Effect (shear) (moment) (shear) (moment) (torsion)

-

.00317 -.000822 Ct = 0 .00188 ex .000605 P .00014 .00004 000002 - .00571 -.00166 -.00099 S + Up and normal to the water plane.

Cyz, + Nose up and about an axis perpendicular to the ~Y' plane of synunetry.

+ To the right and normal to the plane of symmetry.

G..,y, + Nose right and about an axis peTpendicular to ~Z' the water plane.

4rx' + Left wing up and about an axis parallel to the

longitudinal axis.

TABLE E- V • ~ AFT FUSELAGE COEFFICIENTS AT FS 1337.5 FOR 1.20M ANP ~'/ = 67.5° = 1940.0 ft = SAF SREF Q /2 = b /2 820.08 in.

=

AF REF C C 184.05 in.

= =

AF REF Coefficients are based on the air10ads on the fuselage aft of FS 1337.5 and do not include the airloads on the empennage. (Refer to Appendix C for equations which include the air10ads on the empennage.)

C C vy CBY CBZ

vz

~ (shear) (moment) (torslon) Effect (shear) (moment) _.

a. 0 .002173 .00530 a. -.00046 -.000118 ~.0021l -.00030 -.00037 Sa. = 0 c/o -.00007 -.00001 -.00001 c/o Sa.

, 6H .00022 .00006 .00004 6RL .00009 .00001 .00002 .00044 .00011 .00008 P -.00168 -.00034 -.00029 B + Up and nonna1 to the water plane.

CVz, + Aft end up and about an axis perpendicular to the

S3Y'

plane of symmetry.

+ To the right and normal to the plane of symmetry.

+ Aft end right and about an axis perpendicular to the water plane.

'1x, + Left lv-ing up and about an axis parallel to the

longi tudina1 axis.

:rABLE E-VI.- VERTICAL TAIL COEFFICIENTS AT WL 75.0 FOR I.20M AND 1)~ = 67.5 247.4 ft =

Svr

206.76 in.

= bvr/2 188.95 in.

=

-lvr

Coefficients are at the vertical tail root CVTR) and are for use in the equations in Appendix C for the determination of the net airloads coefficients for the aft fuselage point at FS 1337.5.

-

C vy CBX

Crz

Effect (shear) (moment) (tors10n) -.053487 -.023324 .010230 Sa. = Sa. -.001730 -.000755 .000331

a'H

-.002353 -.001621 .001072 ~SP -.000177 .000079 -.000349 00 .009675 .005734 -.006210 .003878 .000352 RL -.001119 P -.003700 -.002645 .002065 R .047596 .022154 -.013930 + To the right and nonna1 to the plane at synnnetry.

Cyy, + Tip to the right and about an axis parallel to ~X' the plane of synnnetry.

Crz' + Leading edge right and about an axi~ perpendicular

to the water plane.

TABLE E-VI!. - LOAD COEFFICIENTS DATA CARD LISTING TABLE E-VII. - Concluded

APPENDIX F RIGID AIRLOAD COEFFICIENTS AT LOAD REFERENCE POINTS. ·CM =0.95) . ' The location of the load reference points where the 'rigid airload coefficients have been determined are presented in figure 'B-1. All of these stations, with the exception of the vertical tail root (VTR), . are the stations where loads will be measured during the B-1 flight loads survey program.

The rigid air load coefficients were determined for each of the applicable aerodynamic effects listed in table I and are presented in tables F-I through F-VI. These coefficients were determined using equations 4, 7, and 10.

The component applicable reference areas, semispans and mean aerodynamic chords are listed in the coefficient tables.

Notes in appendix C which apply to tables B-1 through B-VI for M = 0.85

data apply correspondingly to tables F-I through F-VI for M = 0.95 data.

'fABLE F-I. - WING COEFFICIENTS AT XRS 354 FOR 0.95M AND /\"'J = 67.5°

ft 1946.0 = ~ 820.08 in.

= bw'2 184.05 in.

= <1'l Coefficients are applicable to either left or right wing.

C CTY

Cvz BX

(torsion) Effect (shear) (moment) a 0 .024057 .006690 = -.001643 a .011027 .002776 .000772 • ~ .005623 .001417 .000394 ··.000423 -.000111 .000023 ~SP P ·~.002636 ... 000768 -.000144 .046712 Q .012751 -.000662 ·-.001134 -.000210

o A/S .000139

aa

=

'-.000044 ... 000011

Ba -.000002

A/S .000034 -.000001 .000037 O·Sym Sa = ·-.000081 -.000019 -.000004 Sa Sym + Up and perpendicular to the wing reference plane.

C\rZ'

+ Tip up and about all axis perpendicular to the wing load.reference line (0.36c line)* •.

C ' + Leading edge up and about the wing load reference line

TY (0.36c line)*.

* The wing load reference line passes through the pivot (at XRS 139.515, YRS ·49.845) and the load reference point (at XRS 354, YRS -38.248).

Rev Oct 1981 67 tABLE F-II. ~ HORIZONTAL TAIL COEFFICIENTS AT BP 10.75

EOR 0.9SM AND Aw = 67.5°

= 238.77 ft ~ b /2 = 259.03 in.

m

= 149 .• 38 in.

Srr

Coefficients are applicable to either left or right horizontal tail.

CBX

Cvz Crr

Effect (shear) (moment) (torsion) (). a

= -.183421 -.067341

.041001 ().

.047131 .020727 -.015855 3H .080854 .035558 -.027200

. '

().

.262472 .115430 -.088298 -.020459 -.007027 f3 .002921 3'H .. 021919 .010001 -.006959 (Sym) -.000556 -.000273 .000209 ~SP .000258 .000173 (a/s) -.000169 ~SP -.002591 -.002002 .002197 P Q .551849 .239192 -.239063 Cyz, + Up and perpendicular to the airplane water plane.

C ' + Tip up and about an axis parallel to the longitudinal BX axis.

CrY, + Leading edge up and about an axis perpendicular to the

plane of synnnetry.

TABLE F-III. - VERTICAL TAIL COEFFICIE.WS AT WL 136.56

FOR 0.95M AND f'w = 67.5°

247 .. 4 ft

=

~ 2 206 .. 76 in.

=

hvr/

188 95 in.

=

Cvr

Coefficients are applicable on the upper vertical tail (UVT).

CBX

Cw Crz

, Effect (shear) (moment) (torsion) , 13cx = 0 ... 034024 -.011054 " .001739 -.001155 13cx -.000359 .000056 ., -.007446 -0001977 ~H .000428 ~p -.000252 -.000074 -.000903 ~&RU .013436 .004109 -.004354 :&RL 0 0 P -.003124 -.001146 .000371 R .026671 0008325 -.002780 + To the right and normal to the plane ~f symmetry.

lvY'

+ Tip to the right and ~bout an axis parallel to the ~X' longitudinal axis.

~Z' + Leading edge right and about an axis perpendicular to the water plane.

TABLE P- IV. - fORWARD FUSEL.l\GE COEFFICIR·rrS AT FS 528_5 FOR D.95M AND 1\1 ~ 67.5 ~ ft 5 1946.0 pF ~ 820..0.8 in.

1>pp/2 ~ 184.0.5 in.

~

Coefficients are based on the air10ads on the fuselage forward of FS 528.5.

C C Cyz BY ~ BZ

( ~ )

Effect (shear) (moment) (shear) (moment) torsl0n ~ ex. 0.

.0.0.5325 .00.0111 ex.

.0.0.2413 .0.0.0.685 P .0.0.0.159 .0.0.0.0.42 .0.0.0.0.25 ... 0.0.1424 a -.0.0.4258 - .0.0.0.740.

+ Up and nonnal to the water plane.

+ Nose up and about an axis perpendicular to the plane of syrranetry.

+ To the right and nonnal to the plane of symnetry.

Cw, + Nose right and about an axis peTpendicular to

'13z'

the water plane.

CT.x' + Left wing up and about an axis parallel to the

longitudinal axis.

70.

TABLE F-V. ~ AFI' FUSELAGE COEFFICIENTS AT .

FS 1337.5 FOR 0.95M AND l\~. = 67.5° Z

SAP = SREF· = 1940.0 ft

0AP/Z = b /2 = 820.08 in.

REF = C = 184.05 in.

REF Coefficients are based on the air10ads on the fuselage aft of FS 1337.5 and do not include the airloads on the empennage. (Refer to Appendix C for equations which include the airloads on the empennage.)

.'

CVY' Cvz CBY CBZ

Crx

(shear) (moment) (shear) (moment) (torsion) .Effec.t

a. = 0

.005241 0001425 a. .000020 0000056 .

aa. = 0 c/o

-.001738 -.000251 -.000302

aa. c/o -.000051

-.000007 -.000009 cS' H .000165 .0000434 .0000286 cSRL .000187 .000013 .000032 P .000513 .000126 .000089

a -.000695 -.000141 -.000121

+ Up and normal to the water plane.

+ Aft end up and about an axis perpendicular to the plane of synunetry. ' + To the right and nonna1 to the plane of S)'1:IIlI1etry.

C ' + Aft end right and about an axis perpendicular to BZ the water plane.

Sx' + Left wing up and about an axis parallel to the

longitudinal axis.

:TABLE F-VI. .. VERTICAL TAIL COEFFICIENTS AT WL75.0

FOR 0.9 5M AND '1'/ ;: 67. 50

;: 247.4 ft

Svr

;: 206.76 in.

bVT"2 = 188.95 in.

CVr

Coefficients are at the vertical tail root (VTR) and are for use in the eq~tions in Appendix C for the determination of the net air10ads coefficients for the aft fuselage point at FS1337.S.

C C yy BX Cr~ Effect (shear) (moment) (torslon) -.023234 -.050813

aa = 0 0007872

-.000755 -.001650 .000256 a~ ... 003839

a'H -.003964 .001809

-.000159 aSp 0000039 ... 000346 .008Il2 .013450 .000445 ... 001115 clru.

.004877 -.000274 .... Q02999 .001379 P .018250 -.008767 .039919 R + To the right and noma1 to the plane at symmetry.

Cyy, + Tip to the right and about an axis parallel to ~X' the plane of synunetry.

Crz' + Leading edge right and about an axis perpendicular

to the water plane.

- LOAD COEFFICIENTS DATA CARD LISTING TABLE F-VII.

TABLE F-Vrr. - Concluded

APPENDIX G RIGID AIRLOAD COEFFICIENTS AT LOA}) REFERENCE POINTS (M = 0.70, AW = 67.5°) The location of the load reference pOints where the rigid airload coefficients have been determined are presented in figure B-1. All of these stations, with the exception of the vertical tail root (VTR), are the stations where loads will be measured during the B-1 Flight Loads Survey program.

The rigid airload coefficients were determined for each of the applicable aerodynamic effects listed in table I and are presented in tables G-I through G-VI. These coefficients were determined using equations 4, 7, 10. The component applicable reference areas, semispans and mean aerodynamic chords are listed in the coefficient tables.

Notes in appendix C which apply to tables B-1 through B-VI for M = 0.85 da.ta apply correspondingly to tables G-I through G-VI for M = 0.70 data.

Added Oc:t 1981 TABLE G- I. - WING COEFFICIENTS AT ~S 354 FOR O. 70M AND ~V = 67.5° 1946.0 ft = Sw b /2 820.08 in.

W 184.05 in.

= ~ Coefficients are applicable to either left or right wing.

C C CBX

vz rr

Effect (shear) (moment) (torsion) -.000230 a= 0 .023524 .006454 a .002305 .000802 .009244 .

.009745 .002674 -.000095 a .

-.000313 -.000090 .000015 6SP -.000793 -.000155 P -.00272-5 .051001 .014226 -.000442 Q .000075 J3a = 0 A/S -.000958 -.001894 .0 -.000045 -.000012 J3a A/S -.000212 .000194 J3a = 0 Sym -.000236 .0 .0 .0 J3a Sym ~z' + Up and perpendicular to the wing reference plane.

C ' + Tip up and about an axis perpendicular to the wing load reference line BX (O.36c line)*.

C , + Leading edge up and about the wing load reference line (0.36c 1ine)*.

rr

*The wing load reference line passes through the pivot (at XRS 354, YRS-49.845) and the load reference point (at XRS 354, YRS-38.248).

Added Oct 1981 TABLE G- II. - HORIZONTAL TAIL COEFFICIENTS AT BP 10.75 FOR O. 7 OM and AW == 67. 50 SHT 238.77 ft b /2 259.03 in.

HT SHT = 149.38 in.

Coefficients are applicable to either left or right horizontal tail.

-

C C

c;,z

BX rr (torsion) Effect T) (shea (moment) a .032218 0 910 -.143 -.056550 = .015469 - .011257 a 5424 .03 285 .029383 -.021382 c5H .067 .096446 -.070184 a .220 850 {3 436 -.008 -.003358 .002594 I .04] c5 H .890 .018841 -.012397 c5SP (Sym) -.000 533 -.000268 .000207 c5SP (a/s) .000 122 .000082 -.000075 P -.002 .649 -.001954 .002038 Q .446 420 .212330 -.197940 + Up and perpendicular to the airplane water plane.

+ Tip up and about an axis parallel to the longitudinal axis.

+ Leading edge up and about an axis perpendicular to the plane of synnnetry.

Added Oct 1981 TABLE G-III.- VERTICAL TAIL COEFFICIENTS AT WL 136.56 FOR O. 7 OM AND AW = 67. 50 = 247.4 ft SVT 2 206.76 in.

= bVT/ 188.95 in.

=

CVr

Coefficients are applicable on the upper vertical tail (UVT).

C C

Cyy

BX TZ (torsion) Effect (shear) (moment) {:Jet = -.005120 .004908 0 -.032845 {:Jet -.Ob0089 .000085 -.000569 ·.000857 .000966 c5'H -.006205 c5SP -.000172 -.000025 .000020 c5UR .015006 .002245 -.003670 c5RL .0 .0 .0 -.003054 .000558 .003618 --.003816 R .023071 + To the right and normal to the plane of symmetry.

Cyy, + Tip to the right and about an axis parallel to the longitudinal axis.

C ' + Leading edge right and about an axis perpendicular to the TZ water plane.

Added Oct 1981 TABLE G- IV." FORWARD FUSELAC;E COEFFICIENTS AT FS 528.5 FOR 0.70M AND ~ = 67.5° SFF 1946.0 ft b /2 = 820.08 in.

FF C = 184.05 in.

FF Coefficients are based on the.air1oads. on the fuselage forward of FS 528.5.

C CBZ CTX CVZ CBY vy Effect '(torsion) (shear) (moment) (shear) (moment) Ct = 0 .002696 -.000086 Ct .001980 .00030 .000175 .000023 .000014 I?

f3 ".003466 -.000579 -.000301 Cy-Z' + Up and normal to the water plane.

C ' + Nose up and about an axis perpendicular to the plane of symmetry.

By

~, + To the right and normal to the plane of symmetry.

C ' + Nose r.ight and, about an axis perpendicular to the water plane.

BZ C ' + Left wing up and about an axis parallel to the longitudinal axis.

TX Added Oct 1981 TABLE G-V.- AFT FUSELAGE COEFFIC!ENTS AT FS 1337.5 FOR 0.70M ANDAW = 67.5° = = 1946.0 ft SAF SREF b /2 b /2 820.08 in.

= = REp AF ::: = C 184.05 in.

CAP REF Coefficients are based on the airloads on the fuselage aft of FS1337.5 and do not include the airloads on the empennage. (Refer to Appendix C for equations which include the airloads on the empennage.)

C C CrY

~Z BY CvY BZ

Effect (shear) (moment) (shear) (moment) (torsion) a = 0 .004896 .0007.36 a .000029 .000004 f3a = 0 c/o -.002010 -.000145 -.000175 c/o -.000034 .0 .0 f3a o'H .000722 .000095 .000063 oRL .000229 .000016 .000020 p ,000069 .000565 .000049 {3 - .000672 -.000068 -.000058 ~Z' + Up and normal to the ater plane.

C ' + Aft end up and about an axis perpendicular to the plane of synunetry.

By ~, + To the right and normal to the plane of synunetry.

C ' + Aft end right and about an axis perpendicular to the water plane.

BZ C ' + Left wing up and about an axis parallel to the longitudinal axis.

TX 80 Added Oct 1981 TABLE G-VI.- VERTICAL TAIL COEFFICIENTS AT WL 75.0 FOR 0.70M AND AW = 67.5° 247.4 ft = ~ 206.76 in.

2 = bVT/ ::: 188.95 in.

<=vr

Coefficients are at the vertical tail root (VTR) and are for use in the equations in Appendix C for the determination or the new air10ads coef- ficients for the aft fuselage point at FS 1337.5.

-

C C BX TZ

Sn

(torsion) Effect (shear) (moment)

--

-.003861 {30l = .047194 .010863 {30l -.000067 .000818 .000188 ... 001653 .000789 -.033366 6'H ... 000054 .000013 c5SP -.000236 -.003673 c5Rt] .015024 .004481 -.000300 .004649 .000214 c5RL .0006:51 '-.001028 P -.002981 .035102 .007946 -.003282 R ~ I + To the right and normal to the plane at symmetry.

+ Tip to the right and about an axis parallel to the plane of syrtnnetry.

C ' + Leading edge right and about an axis perpendicular to the TZ water plane.

Added Oct 1981 TABLE G-VII.- LOAD COEFFICIENT DATA,CARD LISTING NASA ARS CASE 4 M=0.70 SW=67.5 WIND TUNNEL DATA RIGID NSFG( I ), SEQuENCE NUMBE~ lXX - WING AT XRS 354.

2XX - I1ORIlONTAL TAIL AT BP 10.75 jXX - VERTICAL TAIL AT WL Ijo.5b ~ __ ~ ____________ 4-"X~X"---_V-,-,E,,,-,-,BllC~L TAl LAT W..L_l~..,lL __ n ___________________ _ ~XX - FORWARD FUSELAGE AT FS 528.:> 6XX - AFT FUSELAGE AT FS 1337.5 CV(I) - COEFFICIENT OF SHEAR EACH EFFECT CB(I) - COEFFICIENT OF BENDING MOMENT EACH EFFECT CT(I) COEFFICIENT OF TORSION EACH EFFECT S,.!:.E~Q~. ~N~O"_'.~_'""'C-.!.V ____ ____"'C._"!.B ___ ___"C'_'_T ______ ,T.uI'__T'__'L ... E~ .. _________ _ 101 0.023524 0.006454 -0.000230 ALPHA = 0 M = .70 SW :: 67.5 102 0.009244 0.002305 0.000802 ALPHA M = .70 SW = 67.5 103 0.009145 0.002674 -0.000095 ALPt1A DOT M :: .70 SW = 67.5 104 -0.000313 -0.000090 0.000015 DELTA SPOILER M :: .70 SW = 67.5 105 -0.002725 -0.000793 -0.000155 ROLL VELOCITY P M = .70 sw :: 67.5 ___ A..1 O!.!!6!L---!0~.~00!..05.!.-'1!W0~0u.lL-,!,!0~.~0A.14:!C2 .... 2!h!6~-~OLl • ....,0",O""O,-"4D4,,,,2----.LP..oLII HJ:L.YELlKlL~ __ ....M....,:: .1.O, __ s.w.. ___ ~...2 107 -0.000958 -0.001894 0.000075 BETA APLHA ZERO A/S M :: .70 Sw :: 07.5 108'-0.000045 -0.000012 0.0 BETA ALPHA AIS M :: .70 SW = 67.5 109 -0.000236 -0.000212 0.(;00194 BETA ALPHA ZERO SYM M = .70SW :: 67.5 110 0.0 0.0 0.0 BETA ALPHA SYM M = .70 SW = 67.5 III 0.0 0.0 0.0 FILLER M = .70 SW = 67.5 112 0.0 0.0 0.0 FILLER M = .m.....s.w :: 67.5 113 0.0 0.0 0.0 FILLER M = .70 SW :: 67.5 114 0.0 0.0 0.0 FILLER M :: .70 SW = 67.5 115 0.0 0.0 0.0 FILLER M =- .70 sw :: 67.5 201 -0.143910 -0.0565,0 0.032218 ALPHA:: 0 M = .70 SW = 61.5 202 0.035424 0.015469 -0.011257 ALPHA M :: .70 sw :: 67.5 203 0.007285 0.029383 -0.021382 DE_LTA H _____ !'l =_..,_1_Q_.s....W. __ ~_b.1,,5 204 0.220850 0.096446 -0.070184 ALPHA OCT M = .70 SW :: 67 .. 5 205 -0.00&436 -0.003358 0.002594 BETA M :: .7e SW :: 67.5

206 0.041890 0.018841 -0.012397 DELTA H PRIME M = .70 sw ; 67.5

207 -o.000!>3;) -0.000268 0.000207 DEL TA SPOILER SVM M:: .70 SW :: 67.5 208 0.000122 0.000062 -0.000075 DELTA SPOILER AIS M:: .70 SW :: 61.5 209 -0.002649 -O.001~!)40.002038 ROLLVfJ.QClTY P M -=---1.Q--.S!t :: 61.5

210 0.466420 0.212330 -0.197940 PITCH VELOCITY ~ M:: .70 SW = 67.5

211 0.0 0.0 0.0 FILLER M ::. .70 SW = 67.5 ZlZ 0.0 0.0 0.0 FIL.bER M :: .70 SW = 67.5

213 0.0 0.0 0.0 FILLER M :: .70 SW = 67.5

214 0.0 o~o 0.0 FILLER .M :: .70 SW ; 67.5 2150.0 0.0 0.0 FILLER M -=------:lD Sw =..b.L..5 301 -0.032845 -0.005120 0.00490B BETA AL~A::O 136.56M:: .70 SW = 67.5 302 -0.000569 -0.000089 0.QOOOS5 BETA ALPHA 136.56M =.70 SW :: 67.5 ---,------ -~--------.-------------------------- :: 303 "'0.006205 -0.000857 0.000966 DELTA H PR IHE 136.56M .70 'sw :: 67.5 304 -0.000172 -0.000025 :: 0.000020 DELTA SPOI LER n6.50M .10 SW :: 67.5 305 0.015006 :: 0.002245 -0.003670 DELTA RUD 136.56M .70 SW .= 67.5 u~ :: 306 0.0 0.0 0.0 DELTA RUD LOW 136.56M .70 SW = 67.5 307 -0.003054 -0.000565 :: 0.000558 ROLL VELO~ P l~Q!!~QM :: gl.5 .1Q S~ 308 :: 0.023071 0.003618 -0.003816 YAW VELOC R 136.56M .70 SW = 67.5 309 0.0 0.0 0.0 FILLER M .70 SW = 67.5 = 0.0 :: 310 0.0 0.0 FILLER M $101;= 67.5 .to 311 0 .. 0 0.0 0.0 FILLER M .70 SW = 67.5

=

312 0.0 0.0 0.0 FILLER M = .70 SW =- 67.5 ___ j)~9 ___ 313 0.0 Q.O F 1 !.I.f K M = .JO... SW = 01.5 314 0.0 0.0 :: 0.0 FIllER M .70 SW = 67.5 315 0.0 0.0 0.0 FILLER M .10 SW =- = 67.5 0.0471910- :: 401 0.010863 -0.003801 BETA AL~A::O WL 75M .70 SW :: 61.5 Added Oct 1981 TABLE G-VII .

Concluded 402 O.OOObl& Q.000lb8 -0.000007 BETA ALPHA

WL 75M .70 SW = 67.5

= 40:; -0.033306 -0.001053 0.000789 DELlA H PR IME SW.:: 61.S WL 75M .70 '" 404 SPQ.lJ.ER :: 61.5 -Q&QQ.f36 -<> ..... .-9.00054 0.000013 DELTA WL 75M .70 SW -,;.~-

=

:: 405 0.015024 0.004481 -0.003613 DELlA RUD UP 75M .70 Slol

WL = 61.5

406 0.004049 0.000214 -0.000300 DEL ·rA

RLO LOW WL 75M .10 S~I = 61.5

=

407 -0.002«;081 -0.001028 0.000631 ROLL VELOC P 15M W!" !!70

S~ = U.2

=

408 0.0:;!:>10Z 0.007946 R

-0.003282 VAW VI::LOC WL 75M .70 S~I = 67.5

=

:: 409 0.0 000 0.0 FILLER M .10 SW '"' 67.5 .~~. ___ 41~..!L~._~_ 0 .. 0 ___ .0..0 FII.t..EB ~ .1Q = Ski '"' Dl.5 :: 411 0.0 0.0 0.0 FILLER M .10 SW :: 61.5 :: 412 0.0 0 .. 0 0.0 FILLER .10 sw '"' 61.5

'" 413 0.0 0.0 0.0 :: Ql.S

F I!.I..E R M .10 sw

=

414 M 0.0 0.0 0.0 FILLER .10 sw '"' 61.5

=

'tl5 0.0 0.0 0.0 FIllER M .10 SW

= 67.5

=

_. ____ . __ ~Jn_(t.Jl.Q.~~?_·L~u'!9..99.9J:~.~ __ ~ • .JL ____ ~.~ P I1A ~..Q...J.Y.f~AL , M_~~ = Dl.S :: 502 0.001'180 O.OOO~OO 0.0 ALPHA ( VERTICAL) M .10 sw

= 67.5

50:; (lATERAl,M :: 61.5 0.000175 0.000023 0.000014 ROLL VEL P .10 Sw

=

:: 504 -0.003 460 -<>.000579 "'0.000:;01 BETA M (kATe RA!. ~ .7Q s~ = Ql.:!

:: 505 0.0 0.0 0.0 FiLLER M .10 Sill :: 67.5 506 0.0 0.0 0.0 IFI LlE R M .70 Sill :: 67.5

=

0.0 0.0 FILLER M :: ~1.S _,~L .... JL • .9.._ !!10 Sill

=

:: 508 0.0 0.0 0.0 FiLLER M .70 Sw

= 67.5

:: 509 0.0 0.0 0.0 FILLER M .10 Sw :: 61.5 a 510 0.0 0.0 0.0 FILLER .1Q s~ M '" DI.:!

0.0 M :: :: 67.5 511 0.0 0.0 FILLER .10 SW :: 512 0.0 0.0 0.0 FILLER M .70 SW :: 67.5 :: 0.0 M

.. __ ~.L.~_~_._9._.JL_ .. _ FIlleR .1Q sw

'" Dl.S :: 514 0.0 0.0 0.0 FILLER M .10 sw .. 61.S 0.0 515 0.0 0.0 FILLER .10 SW = '" 61.5

'" -_._-_._-

.- .. -~-.-.- ... - .. -- .-'-._ .. _ ..

001 +0.004 b96 0.OO073() 0.0 (VERTICAL) ALPHA=O M .70 SW

= 67.5

=

602 o .OQ0029 0.OQOOO4 O Oe ALPHA .IQ I ~I:R II tA L » M = S~ = D1.5 603 :: -0.002010 -0.0001 .. 5 -0.000115 BfH AL PHA *0 C/O LATM .70 SW = 61.5 604 -0.000034 0.0 0.0 ALPHA (LAT)M

BETA C/O .70 Sw = 67.5

=

605 0.000722 :: O. Q.QQQj,'2 Q!QQQQQl H PR H!E

QHIA IL.A!! M .10 SW = b.1.5

606 0.000229 O~OOOOlb 0.000020 DE:LTA ROO LOW (LAn M :I: 67.5 .70 SW

=

607 0.000565 0.000069 O. (JOO049 ROLL VELOCITY LAT M .70 P SW = 67.5

=

608 -0.000072 .-0.000068 -o.OOOO~8 B~TA i !. AT I: RAI.l

M .10 Sw = Dl.,5

=

609 0.0 0.0 0.0 FillER M .10 SW

= 67.5

=

610 0.0 0.0 o.d FILLER M .10 SW = 67.5 = 611 o .Q 0.0 0,0 E lLL.E B M .10 :: 01.5 Sill

=

612 Cl.O 0.0 0.0 FILLER M .70 SW

= 67.5

=

0.0 613 c.O 0.0 FILLER M .10 = Sw = 67.5 614 0.0 0.0 . 0.0 FI!.!..I:R M .1Q Sit :: 61.:!

=

615 0.0 0.0 0.0 FILLER M .70 sw =. 67.5

=

Added Oct 1981

APPENDIX H RIGID AIRLOAD COEFFICIENTS AT LOAD REFERENCE POINTS (M == 1.60, AW = 67.5°) The location of the load reference points where the rigid airload coefficients have been determined are presented in figure B-1. All of these stations, with the exception of the vertical tail root (VTR), are the stations where loads will be measured during the B-1 Flight Loads Survey program.

The rigid air load coefficients were determined for each of the applicable aerodynamic effects listed in table I and are presented in tables H-I through H-VI. These coefficients were determined using equations 4, 7, and 10. The component applicable reference areas, semispans and mean aerodynamic chords are listed in the coefficient tables.

Notes in appendix C which apply to tables B-1 through B-VI for M = 0.85 data apply correspondingly to tables H-1 through H-V1 for M = 1.60 data.

Added Oct 1981 TABLE H-I.- WING COEFFICIENTS AT XRS 354 FOR 1.60M AND~ = 67.5° ~ = 1946.0 ft b /2 = 820.08 in.

W Sw = 184.05 in.

Coefficients are applicable to either left or right wing.

-

CBX elY CVZ Effect (shear) (moment) (torsion) (lr = 0 .020939 .006560 -.002288 cr .010702 .003193 .000306 c:r .0 .0 .0 CiSP .000020 -.000148 -.000030 P -.002472 -.000757 -.000063 Q .035134 .010720 -.002668 .000291 f~a = 0 A/S -.000619 -.000151 -.000454 -.000115 -.000038 13a A/S 13& = 0 Sym .000045 .0 -.000135 .000025 .000025 -.000051 13a Sym CyZ' + Up and perpendicular to the wing reference plane.

C ' + Tip up and about an axis perpendicular to the wing load reference line BX (0.36c line)*.

CIT' + Leading edge up and about the wing load reference line (0.36c line)*.

*The wing load reference line passes through the pivot (at XRS 139.515, XRS-49.845) and the load reference point (at XRS 354, YRS-38.248).

Added Oct 1981 85 TABLE H~II.~ HORIZONTAL TAIL COEFFICIENTS AT BP 10.75 FOR 1.60H AND AW = 67. 50 SHT =·238.77 ft b /2 = 259.03 in.

HT C = 149.38 in.

HT Coefficients are applicable to either left or right horizontal tail.

C Crr

CVz

BX Effect (shear) (moment) (torsion) a = 0 -.082380 -.028675 .029277 O!

.020148 .009443 ~.010049 dH .038529 .017524 -.019050 IY. .119250 .061051 -.063703 ~ -.020150 - .007246 .003210 c5'H .017256 .007993 -.008290 c5SP (Sym) -.000386 -.000168 .000195 c5SP (a/s) .000147 .000077 -.000090 P -.002359 -.001922 .002119 Q .393370 .183300 -.219810 + Up and perpendicular to the airplane water plane.

CVZ'

+ Tip up and about an axis parallel to the longitudinal axis.

+ Leading edge up and about an axis perpendicular to the plane of synnnetry.

Added Oct 1981 TABLE H-III.- VERTICAL TAIL COEFFICIENTS AT WL 136.56 FOR 1. 60M AND AW = 67.5° 247.4 ft = SVT 2 = 206.76 in.

bVT/ 188.95 in.

=

CVr

Coefficients are applicable on the upper vertical tail (UVf).

C C Cyy TZ BX (torsion) Effect (she ar) (moment) {310! .006545 0 -.03 4658 -.005487 = (310!

-.000151 .000180 -.00 0953 v 6 H 2820 -.000123 .000583 -.00 oSP -.00 0122 -.000020 .000020 oUR .000825 .001954 .00 5500 oRL .0 .0 .0 P .001090 -.00 4395 -.000856 7.1527 R .OJ .005128 -.006678 + To the right and normal to the plane of symmetry.

+ Tip to the right and about an axis parallel to the longitudinal axis.

.)- C Leading edge right mld about an axis pe~endicular to the TZ' water plane.

Added Oct 1981 TABLE H~IV.- FORWARD FUSELAGE COEFFICIENTS AT FS 528.$ FOR 1.60M ANDAW = 67.5° ;:; 1946.0 ft SpF b /2 = 820.08 in.

Fp ;:; 184.05 C in.

FF Coefficients are based on the airloads on the fuselage forward of FS 528.5.

C e

CBY vy GTX

Cvz BZ

Effect (shear) (moment) (shear) (moment) (torsion) & ::c 0 .000481 -.001118 & .001812 .000256 p .001221 .000088 !000106 -.004710 -.000340 -.000409 CVx' + Up and normal to the water plane.

CBY' + Nose up and about an axis perpendicular to the pl~e of symmetry.

~, + To the right and normal to the plane of symmetry.

C ' + Nose right and about an axis perpendicular to the water plane.

BZ C ' + Left wing up and about an axis parallel to the longitudinal axis, TX Adcied Oct 1981 TABLE H-V. AFT FUSELAGE COEFFICIENTS AT FS 1337.S FOR 1.60M andAW ~ 67.5° SAF = SREF = 1946.0 ft b /2 = b /2 = 820.08 in.

REF AF C 184.05 in.

REF Coefficients are based on the airloads on the fuselage aft of FS1337.5 and do not include the air10ads on the empennage. (Refer to Appendix C for equations which include the airloads on the empennage.)

C C C C BY Vi BZ TY

s'z

(torsion) (moment) (moment) (shear) Effect (shear) .001981 a = 0 .005458 a .0 .0 -.000200 -.002300 -.000166 (3a = 0 c/o .0 .

-.000070 .0 (3a c/o -.000070 -.000803 -.000105 6'H .0 .000018 .0 6RL p .000030 .000340 ,000042 (3 -.000102 -.001170 -.000119 C ' + Up and normal to the water plane.

vz

C ' + Aft end up and about an axis perpendicular to the plane of symmetry.

By C , + To the right and normal to the plane of symmetry.

vY

C ' + Aft end right and about an axis perpendicular to the water plane.

BZ C ' + Left wing up and about an axis parallel to the longitudinal axis.

TX Added Oct: 1981 TABLE H-VI.- VERTICAL TAIL COEFFICIENTS AT WL 75.0 FOR 1.60M and AW = 67.5° ~ = 247.4 ft bVT/2 = 206.76 in.

= 188.95 in.

';' Coefficients are at the vertical tail root QrrR) and are for use in the equations in Appendix C for the determination of the new airloads coefficients for the aft fuselage point at FS 1337.5.

C C

Cyy

BX TZ (torsion) Effect (shear) (moment) {3a .011700 -.006195 0 .053847 = -.000322 .000170 {3a. -.001471 o 'H -.002021 -.000480 .000468 oSP -.000157 -.000040 .000018 .001645 -.001956 oRU .005506 -.000521 c5RL .002729 .000121 P -.003603 -.001475 .001220 .011025 - .007172 R .047887 + To the right and normal to the plane at symmetry.

+ Tip to the right and about an axis parallel to the plane of symmetry .

C ' + Leading edge right and about an axis perpendicular to the TZ water plane.

Added Oct 1981 TABLE H-VII.- LOAD COEFFICIENTS DATA CARD LISTING NASA ARS CASE ~ M:.l.e SW=67.5 WINO TUNNEL DATA RIGID _~~~f ~J LL.~;"Q!,l.f-".J!.~ _~!,J~~ ~B ... _.-!.!~_ - WINGo A T...Mll.~ ____ . __ .

2XX - t-()RllONTAl TAll AT BP 10.7~ 3XX - VERTICAL TAXl AT I'll 136.50 4XX - VERTICAL TAll,. AT loll,. 75.0 5XX - FDRWARU ~U~ElAGE AT F~ 528.5 bXX - AFT FUSELAGE A1 F50 13:n.5 ._~YLU __ :: .. ~P~Ef.J-':'_HNJQf S!"1M!L~ CH __ J:.£f.r;~l_ .. _____ .. _ .... _ ...... _. ___ . __ . ________ _ Col I I - COEFFICJF.Nl OF BENDING MOMENT EACH EFFEC T CTI II - C8EFFICIENT OF TORSION EACH EFFECT S~Q •. N...CI. c:.V. CB .{'L.. TI1I..1: 101 0.020':139 0.000500 -0.OO'::2tH! ALPHA:. 0 M = 1.0 SW = 67.5

102 0.010102 O.0031~~ 0.000306 ALPhA M = 1.6 sw ::. 67.5

103 0.0 0.0 0.0 . ALPH~_JH~L_. ___ u_ M :; 1.6 SW ::. 67.5

~-.--- 104·-=o:o·oo·i4q··'.:o-~oooo2:0· O.OOOOlO DELTA SPOILER M = 1.0 sw ::. 67.5

105 -0.002472 -Q.000757 -O.Ouv063 ROLL "ELOCllYP M = 1.6 SW ;: 67.5 lQ~ u.V35134 0.01(.1720 -Q.(,JOZQol) PITCH VELOCITY Io.i. M.. =. 1.b SW :. 67.5 107 -O.OOObl'.! -0.000151 0.0002<tl BETA AlPHA=O A/~ M = 1.0 Sw = 07.5 106 -0.00045 .. -o.OOOllS -0.000(.';;8 BE.TA ALPHA AIS M ::. 1.0 sw = 67.5 . __ 1 O~_.q.~ OQ!'J..!:~.:L __ Q!. Q.. __ . _._._-:Q~..l~5 81: T UL PH" =0 .s..y M ... _ M = ... J ~~....sJL....:;_~.1~ 110 O.OOO(;2~ O.OOOOZ!) -0.000051 BETA ALPHA SYM M ;: 1.0 SW ::. b7.5 111 0 • 0 U. CO. C r 11.. LE R M :. 1.0 S 1'1 ::. 67.5 U4 Q~(.I (,J~(., .. Q.(I f.lLJ"EiL M = 1.b SW = 67.5 113 o.a 0.0 0.0 FILLER M ::; 1.0 SW ::; 07.5 114 0.0 0.0 0.0 FILL,ER M :. 1.b SW :; 67.5 .. _. __ .. U5. ...... Q!~._._ ... _ .. 9 !9_._ ....... 9 . .!.o_._._,_._fl!,J,.,-I;.~_ __ . ... t'(.::: .1. •. 9 s..W .. _=.fl.7~.2 ?01 -V.O!lZ~8(1 -Q.02B67!) 0.02'1217 ALPHA::. 0 M ::. 1.0 Sill :; 07.5 Z02 0.(.;2(;146 O.CO':l44~ -O.01U04'l1 ALPhA M ;: 1.0 Sid ::; b7.5 203 0.031>529 O.01152't -o.uhI,J50 llELTA H M :. 1.0 sw :. 67.5 204 0.11"',50 0.001051 -o.Ob:n03 ALPHA pUT M '" 1.6 sw ::. b7.5 20~ -0.020150 -0.007240 0.OO~210 BET~ M = l.b SW :. 67.~ 20b .. Q.9.1"T?So Q.\lQ79~~~.Q .•. YQ.~~':I9.Q.f.L T1LJ:L.e.8.1.ML._. _ .. _.f:\. .1~Q. S)1,:;o67.:;' 207 -Q.l.t,u:'8b -Q.OOOl08 0.00(;195 DELTA SPOILER S'IM M:; 1.0 Sw ::: b7.5 208 0.000147 0.000077 -O.OOvO"'O DlLTA SPOILER A/S M '" 1.b SW ::; 67.5

209 -0.0\)2:'5<'- -0.0(11'122 0.002119 ~UI.I" VELOCITY f' M :; l.t. SW = 67.5

210 U.i>93370 0.183300 -O.'l'1lilO PITCH "ELOCITY '" M:; l.b SW ::; 07.5 211 0.0 0.0 O.C! FILLER tI. ::; 1.6 50W = 67.5 21? (J.U 0.0 O.Q fH.l.EJL_.... . ..... 11....."" 1. •. o .. SW .. E .. ~J .• ~ 2 13 (0 • 0 0 • 0 0 • u F I I.. LE R M ::. 1 • 0 S \oj = b 7 • ~ 214 0.0 0.0 0.0 FILLE;R M = 1.0 SW ::; 67.5 2 15 0 • () 0 • 0 0 • 0 F I L LE R M = 1. b Sw :. 67.5

301 -0.034tt5h -0.00:;4,>7 0.000545 tlETA ALPHA=O B6.50M = 1.0 SW = 67.5

302 -O.00C~~3 -o.000l~1 O.OOOlbO BETA ALPhA 1~6.~6M ::. 1.6 Sw ::. 67.5 303 -C~C0282C -,).000123 0.00(.;583 Dt;.LTA H PRIME 1.36.56"'1 = 1.6 SW : 67.5 :3 04 -0. CiuO 122 . :-:-:{I._999.Q,29 . .,. Q •. QQ.iJQ.~l(_Df.;1. T~ ... ~.£QU ~JLl.;,.Q .•. 2.Q.M._.= .. .l_ .. 6._SV.L :;._~ 1.S -- - 305--0-:-005:'6'<:; 0.000825 0.v019~" DELTA RUD UP Uo.50M = 1.b sw ::: 67.5 306 0.0 0.0 0.0 DtLlA RLO LUW 1.:>6.56M ::. 1.6 SW :: 67.:; 307 -0.004~95 -Q.00085c1 0.001090 RUI..L VE-LtJC P 136.56M::. 1.6 Sw :: 67.5 308 O.O~I~21 0.00512u -0.000078 PITCH VELOC Q 13o.So~ = 1.0 Sw :: 07.5 309 0 • Ci 0.0 0 • 0 F I L LE R M = 1.6 50 1'1 = 07.5 __ ~_~_!JL_O...!..Q ... __ ....... Q~_9 ... ___ ._ . ..9_~V~ ... _ .. ___ fl\..I.r;.K, .1'1 ... _: ... 1..0. ... SW .. = .. 61.5 311 0.0 0.0 0.0 FIl.LER M :; 1.0 SW :. 07.5 312 0.0 0.0 0.0 FILU;R M = 1.6 51'1 = 67.':> 31.3 O.C 0.0 O.lI FIl..LER ,.. :: 1.6 SW ::; 61.5 314 (j.G v.D O.lI FILl,Ef< M = 1.0 SW :: 07.5 315 Q.O 0.0 0.0 FILU:K ,.., = 1.6 SW - 07.5 401 0.053 .. 87 0.0111UO -0.OUo.l95 BUA ALf'HA=(; wL 751'1 = 1.0 .. Sw : 07.5 Added Oct 1981 91 TABLE H-VII . Concluded -- --------_. __ ._------ --"" ;::: 402 -0.v01,+71 -o.0vv322 0.000170 BETA ALPHA WL 75M = 1.6 sw 67.,:> -0.002021 -0.000480 ;::: 403 0.00ll400 DEUA h PR IME WL 7':>1'1 1.0 Sw 07.':> :;. ;::: . __ 4Q4 -0 .• 000157 -0.000040 O. 0(HlO 18 OlLTA 5POI LEk WL '15M 1.6 Sw 67.~ ;::: 405 0.005 ':>00 0.001645 -0.001956 DELTA RUO UP WL 75M 1.b SW = 67.5 ;::: 406 0.002"12'1 0.000121 -0 .• 000521 01:: L TA RlJD LOW WI. 75M 1.6 sw 07.5 = -0.003_003 -a.OOh-/5 407 0.00122(.; RULL VE J..ill;_L. _.~5.M..~....l._.1L~,"--..b.lL2 ;::: ;::: 408 0.0471"87 .0.011025 -0.007172· YAW VELDC R WI. 75M 1.6 SW 67.5 ;::: 409 0.0 0.0 0.0 F ILLE R M 1.6 Sw = 67.5 . _ .. 41.9 __ 0.LQ ___ _Q .• Q_ .... 0 •. 0_ . F 1 LL.E.R ..

--_ ... .... _ ......... M _=_.l.lL Sw ... =- .. 0.7.5 ;::: ;::: 411 0.0 0.0 0.0 F 1 L LE r< M 1.6 sw 67.,:> ;::: ;::: 412 0.0 0.0 0.0 FIL LE k M /.)7.5 1.6 SW 61 __ 5 413 0.0 __ ~&.... __ ~~ ___ -..EJ~~R __ . ___ . __ . ___ M ... ~ .. l . ..p .. .s.\L.~ ;::: :: 414 0.0 0.0 0.0 F 1 L LE R M 1.6 Sw 07.5 ;::: :; 415 0.0 0.0 0.0 F 1 LLE R M 1.6 Sw 67.5 ;::: 50! 9!000481_ ~G. 00 11 lti 0.0 ALPHA ,,:0 ( V E81I (;ALI .. M. = 1.6 Sw 67.5 ".

;::: ;::: 502 0.001tl12 0.0002'>6 0.0 ALPhA ( VERTICAL) M 1.6 SW 67.5 :: 503 o .0012d 0.00006b 0.OCl.ilO6 BlTA ALPHA=O CIO LATM = 1. b sw 67.5 ;::: 504 -0.004"10 -0.0003,+0 -0.00040'1 BETA (LATERAL) M = 1~6 s.w 67.5 ;::: ;::: 505 0.0 o.v 0.0 F I LLE R M 1.6 Sw 67.5 ;::: 0.0 506 0.0 0.0 FILLEK M "- 1.6 Sw 67.5 O.G F lLI,.ER = Sw .-'?J2.1_...Q.~_ .9.~0 M .. 1.6. =- 67.5 ..

;::: 508 0.0 :: 0.0 0.0 FILLE R M 1.6 sw 07.5 ::.

509 0.0 0.0 0.0 F I L LE R M 1 .. 6 67.5 =- SW 0.0 ::.

510 0.0 0.0 FI L LE R M 1 .. 6 sw 67 •. 2.

= ::.

511 0.0 0.0 0.0 FILLER M 1.6 sw 07.5 = ;::: 0.0 0.0 0.0 FI LLE k 512 M ='1.6 SW 67.':> ::.

0.0 0.0 FILLE R M . -2..1 ~._ ...9~_Q_.. _ 1.6 sw Q 7.5 =- ...

;::: 0.0 ::.

514 0.0 0.0 F ILLE R M 1.6 SW 07.5 ,0.0 ::. ::.

515 0.0 0.0 FILLER M 1.6 SW 67.5 .-_._--"-"-----------_ .. _-, ....-.-- 601 0.005 .. 5& 0.0019131 0.0 ALPhA=O (VERTICAL) M = 1.6 Sw ;::: 67.5 __ .. __ .. Q.QL .Q~~.___._ . Q .• 0 .o .• JL .. .t\bPbl,---LVf..!LT.llA.l...L ___ ~_-'E .. ..l..JL.S.\i._::_ilL2 603 -O.0u2::'Ov -{,.OOOI60 -0.0002(;0 SUA ALPHA;:::O CIO LATM ::. 1.0 SW ::. 07.5 604 -0.(,00(;70 O.LI 0.0 BI:TA ALPHA C/O (LAT)M ::. 1.6 sw :; 67.5

605 -O.OvOt{Oj -0.000105 -o.Ouvo·,o DI::L1A h PRIME (LAT) M ::. 1.6 SW = 67 .• 5

606 0.000ull:l 0.0 o.() DH1A ROO LOW (LAT) M = 1.0 Sw : 67.5 607 0.000340 0.0000'02 0.000030 ROLL VI:LOCI1Y PLAT M ::. 1.6 Sw ::. 67.5 ooa -O.OOU.7..0 -9 ..•. 0'O.9.1..A.'t .. ~q .•. J,J()VJ()~ .~...l;.I!' ..... 1.l..61.E . .!(_~.J....L ____ ~= 1.6 SW ::. 67.5 609 0.0 0.0 O.l! FILLEt< M :: 1.0 SW ::. 07.5 610 O.V C.O 0.0 FILLER M = 1.6 SW ;::: 67.5 611 v.v 0.0 I).V FILLEK M ::. 1.6 sw = 67.5 612 0.0 0.0 0.0 FILLI:R 1"1 = 1.0 SW =- 67.5 613 0.0 0.0 0.0 FILLEk M = 1.6 Sw ;::: 67.5 (:) 14. .9.0 u. 0 .. 0." Q ...Ft!.. L,i;~ .... _ .. _ ...... _ ._ .... J~_~_lLQ...ft._~_2.1..2 615 O.lI 0.0 0.0 FILLER M = 1.6 Sri = 67.,:> Added Oct 1981

APPENDIX J RIGID AIRLOAD COEFFICIENTS AT LOAD REFERENCE POINTS (M = 2.20, AW = 67.5°) The location of the load reference points where the rigid airload coefficients have been detennined are presented in figure B-1. All of these stations" with the exception of the vertical tail root (VIR), are the stations where loads will be measured during the B-1 Flight Loads Survey program.

The rigid air load coefficients were determined for each of the applicable aerodynamic effects listed in ta.ble I and are presented in tables J-I through J··VI. TIlese coefficients were detennined using equations 4, 7, and 10. The component applicable reference a.reas, semispans and mean aerodynamic chords are listed in the coefficient tables.

Notes in appendix C which apply to tables B-1 through B-VI for M = 0.85 data apply correspondingly to tables J-I through J-VI for M = 2.20 data.

Added Oct 1981 TABLE J-I.- WING COEFFICIENTS AT X 354 FOR 2.20M ANDAw = 67.5° RS Sw = 1946.0 ft b /2 = 820.08 in.

W ~ = 184.05 in.

Coefficients are applicable to either left or right wing.

C ClY Cvz BX Effect (shear) (moment) (torsion) & = 0 .018800 .005150 ~.003918 & .009300 .002617 -.000384 0/ .0 .0 .0 ..

6SP -.000097 ~.000020 .000016 -.000036 P -.002083 -.000648 .025344 .008402 -.003291 Q -.000127 .000291

13& = o A/S -.000743

j3& A/S .000055 .000011 .0 -.000116 -.000036 ~& = 0 Sym -.000164 13& Sym -.000179 -.000054 .000058 CvZ' + Up and perpendicular to the wing reference plane.

C ' + Tip up and about an axis perpendicular to the wing load reference line BX (0.36c line) *.

ClY' + Leading edge up and about the wing load reference line (O.36c line)*.

*The wing load reference line passes through the pivot (at XRS 139.515, YRS-49.845) and the load reference point (at XRS 354, YRS-38.248).

Added bct 1981 TABLE J-II.- HORIZONTAL TAIL COEFFICIENTS AT BP 10.75 FOR 2.20M AND AW = 67.5° 238.77 ft = SHf 259.03 in.

= b l2 H 149.38 in.

= CHf Coefficients are applicable to either left or right hOFizonta1 tail.

C

;X

CVz TY

(moment) (torsion) Effect (shear) a -'.020695 .022117 '- 0 -.062590 a .018928 .009394 -.010369 oH .025820 .011969 -.012889 .

a .048821 .024229 -.026745 {3 .0 .0 .0 -.011501 o'H .023039 .010680 ... 000039 .000045 OSP (Sym) -.000089 -.000020 oSP (a/s) .000034 .000018 P .001867 -.002071 '-.001590 -.144600 .254600 .111350 Q + Up and perpendicular to the airplane water plane.

CVZ'

+ Tip up and about an axis parallel to th~ longitudinal axis.

+ Leading edge up and about an axis perpendicular to the plane of symnetry.

Added Oct 1981 TABLE J-III.- VERTICAL TAIL COEFFICIENTS AT WL 136.56 FOR 2. 20M AND AW = 67.. S° 247.4 ft = ~ 2 206.76 in.

= bVT/ 188.95 in.

=

CVr

Coefficients are applicable on the upper vertical tail (UVT).

C C BX TZ ~ Effect (shear) (moment) (torsion) {Ja -.023927 -.003838 .004513 = 0 {Ja -.000489 -.000078 .000092 .000386

o 'H -.OO18~5 -.000081

o SP '-.000015 .000015

-.000091

oUR .003299 .000499 -.001192

oRL .0 .0 .0

P -.003896 -.000788 .000983 .002966 -.004283 R .018237 + To the right and normal to the plane of symmetry.

~, + Tip to the tight and about an axis parallel to the longitudinal axis.

Leading edge right and about an axis perpendicular to the water plane.

Added Oct 1981 TABLE J - IV. - FORWARD FUSELAG'E COEFFICIENTS AT FS 528.5 FOR 2.20M ANDAW = 67.5° SFF = 1946.0 ft b /2 = 820.08 in.

Fp C ::: 184.05 in.

pF Coefficients are based on the airloads on the fuselage fonvard of FS 528.5.

C C C CVZ CvY TX BY BZ Effect (shear) (moment) (shear) (moment) (torsion) 0' = 0 -.003902 -.001250 0' .001489 .000233 .000056 .0 .0 P -.006606 -.001104 -.000574 13· -- Cy-Z' + Up and normal to the water plane.

C ' + Nose up and about an axis perpeJ?dicular to the plane of synnnetry.

By lvY' + To the right and nomal to the plane of symmetry.

C ' + Nose right and about an axis perpendicular to the water plane.

BZ C ' + Left wing up and about an axis parallel to the longitudinal axis.

TX Added Oct 1981 TABLE J -V.- AFT FUSELAGE COEFFICIENTS AT FS 1337.5 FOR 2.20M andA = 67.5° W SAP = SREF = 1946.0 ft b /2 = b /2 = 820.08 in.

AF REF = C = 184.05 in.

REF Coefficients are based on the airloads on the fuselage aft of FS1337.5 and do not include the airloads on the empennage. (Refer to Appendix C for equations which include the airloads on the empennage.)

C C C c;,Z BY ~ BZ TY (shear) (moment) (torsion) Effect (shear) (moment) a = 0 .008028 .001533 a .000142 .000008 {3a ::::

o c/o -.000104 -.000014 .0

-.000025 .0 .0 {3a c/o , .000024 .0 .0 o'H .000041 .0 .0 6RL P .000179 .000022 .000016 -.001422 -.000144 -.000124 {3 c;,Z' + Up and normal to the water plane.

C ' + Aft end up and about an axis perpendicular to the plane of symmetry.

By ~, + To the right and normal to the plane of symmetry.

C ' + Aft end right and about an axis perpendicular to the water plane.

BZ C ' + Left wing up and about an axis parallel to the longitudinal axis.

TX 98 Added Oct 1981 TABLE J-VI.- VERTICAL TAIL COEFFICIENTS AT WL 75.0 FOR 2.20M ANDAW = 67.5° = 247.4 ft

\rr

2 = 206.76 in.

bvrl = 188.95 in.

~ Coefficients are at the vertical tail root (VTR) and are for use in the equations in Appendix C for the determination of the new air10ads coefficients for the aft fuselage point at FS 1337.5.

C C TZ

Sri BX

(moment) (torsion) Effect (shear) .008037 -.004340 l3a 0 .035396

=

-.000164 .000089 -.000724 l3a -.000317 .000310 <5'H -.001337 -.000030 .000013 c5 S1' -.000118 .000990 -.001193 c5RU .003302 .000889 .000039 -.000174 c5RL p -.001323 -.001025 -.002998 .029111 .006471 -.004185 R C ' + To the right and normal to the plane at synnnetry.

vy C ' + Tip to the right and about an axis parallel to the plane of BX synnnetry.

C ' + Leading edge right and about an axis perpendicular to the TZ water plane.

Added Oct 1981 TABLE J-VII.- LOAD COEFFICIENTS DATA CARD LISTING ~ASA A~S CASE 0 M;2.~C SW=o7.5 WINU TUNNEL DATA RIGID NSF;:; I I) SEQI,JENCJ; __ ,..UMB'=K .1XX - WiNG AI XR~ ____ . ______ _ 2XX - ~OKIZONTAL TAIL AT ~P 10.75 3XX - VEKTICAL TAIL AT WL Do.56 "XX - VERI leAL TAl L I>T WI. 75.0 !)XX - F-UkWARD FUSELAGE AT FS 528.5 oXX - AFT FUSELAGE AT FS 1337.5 _~ ~~\lJIL:: ___ C1LE.fEl_~,JJ:J~L9.£_-.-S!:1E ~K __ ~"-tl __ . ~Efe_c..l _______ ..

Call) - COtffltlE~T OF BENDING MO~ENT EACh EFFECT CT(l) - COEFFICIENT OF TORSION tACH EFFECT SE'-I.NQ. tv CQ_. __ .J.L _ I1.TLE.

101 O.OlbtOO 0.005150 -0.003'118 ALPhA=- 0 M ::. 2.2 SW = 67.5 102 O.009~0(; 0.002017 -0.{J003b4 ALPMA M = 2.2 SW = 67.5 _________ 103_---.2.~___ 9.0 0.0 AL._e.HA_DJ::[L________ M ::.~-.--S1L::. 07.5

104 -0.000097 -0.000020 O.GOu016 DELIA SPOILER M = 2.2 SW = 67.5

105 -0.002083 -0.0000,+8 -0.000030 ROLL VELOCITY PM::. 2.2 Sw ::. 67.5

106 0.Ol5.)4" 0.008402 -0.00.>291 PITCH VELOCITY r.. _M';. 2 .. 2. .. SW = __ 6.1.5

107 -0.000143 -0.000127 0.000290 BETA ALPHA ZERU A/S M ::. 2.2 Sw = 67.5 108- 0.000055 O.OOOOll 0.0 BETA ALPHA AIS M :I 2.2 sw = 61.5 ____ ~O_9 __ ::_Q_!_Q(,I.9_'_iQ.~_:Q.000UJL_~0. 000036. BET A ALPHA ZER_Q_~Y1Llt~2..2_._S.w~1___.5..

110 -O.(J00179 -0.000054 0.uOo05& BI:TA ALPHA SYM M = 2.2 SW ::. 67.5

111 0.0 0.0 0.0 FILLER M = 2.2 sw = 07.5 112 o~o _____ Q.O___ _____ Q._Q__FILLE_R .1". ... ,.2. SW '=_.07 .. 5_

113 0 .0 0.0 o. 0 F IL LE R M = 2.2 S W :: 67.5

114 0.0 0.0 0.0 FILLER M = 2.2 SW = 07.5

___ .. li~___.9 __ ._.9 ______ Q~_9_· ____ ~ FILLER ,., = 2.2 sw ::. 67.5

201 -0.Ob2590 -0.020095 0.02.d11 ALPHA:: 0 M = 2.2 SW = 67.5

202 0.018"21:1 0.0(',93';14~. OlO~b9 ALPHA M :: 2.2 sw ::. b7.5 .. 4{L3 __ 0~02!) __ B~_9 0.011969~Q~OJ.2ttfl9_0ELlA_bM_:;: 2.2SW:; 07.5 204 0.048821 O.0~4229 -0.026745 ALPHA DOT M = 2.2 SW = 67.5 205 0.0 v.o 0.0 BETA M :: 2.2 SW :: 67.5 206 0.0230:;9 0.010680 -0.011501 DI::LTA H PRI~ M :I 2.2 Sw ;:: 67.5 207 -o.00008~ -0.000039 0.000045 DELTA SPOILER SVM M:: 2.2 SW = 67.5 208 0.V00034 0.000018 -0. v(J0020 DEllA SPOI LEk A/S M:: 2.2 SW = 67.5 .. _______ 4. Q~L::IL~Q 02_ 9VL __ -::Q ~_QQ155<t ___ Q_._1I(LJ,a_61._BULL_\'.E LQC_l1~..P_. ___ .. __ . ___ .~L.:::: ._2.2_ s.w.. _.~ ... b7_. 5_

210 0.254000 0.111350 -0.144000 PITCH VELOCITY ~ M = 2.2 SW = 67.5

2 11 0 .0 0 • 0 0 • 0 F I L LE R M 2 .2 S W = b 7 • 5

212 0.0 0.0 0.0 ElLLER M '" 2.2 SW = 67.5

213 0.0 0.0 0.0 FILLER M = 2.2 sw = 67.5

2 14 0 .0 o. 0 0 • 0 F I L LE R M :: 2.2 S W = 07. 5 .. __ ~ 1 ~ ___ . Q ~JL _____ .. _O~Q_ .. ________ .0_. __ 0.___ . __ E..lLLE. K ___ ._ ... ~ _____ ... __ ... __ .... M .. :: _2...2. __ ..s.~L:Lb1_ ... 5.

301 -0.023"27 -0.003838 0.u04513 BETA AL~A=O 1~6.50M = 2.2 Sw = 67.5

302 -0.000489 -0.000078 o. v00092 BETA ALPHA 1~6.56M = 2.2 SW :: 07.5

303 -0.v()lb6~ -0.000081 O.OOU3bt> Dt:LTA H PRIME Bo.50M ::. 2.2 Sw = 67.5 __________ }04: __ ,:,,9 !. CiQ.Q ~'.1J __ ::9!. Q.9.Q9J'? ___ O_~_Q.9 00 15._0J:_L __ 1!\_~. POJ1.J;_~_._l~~_ • .2.9~_"':._Z ~LS.!L --=--_9-.1...-.2 305 o .OO32'~9 0.0004'19 -O.OOll92 DI::LTA RLO UP 1~6.56M = 2.2 sw 67.5 = 306 o .0 0.0 0.0 ::: DEllA RUO LOw 1~0.50M 2.2 Sw 07.5 = 307 -0_. 003 b~ t?

-o.OOO7bi) ... O._QO_U-"&3 ROLl, VELDC P l;:)() .56M._::: __ 2~2 .. sw .::: ... 67.5 308 0.0 it!:::3 7 0.002900 -0.OO .. 2b3 'fAW VHOC R Bb.56M ::. :0 2.2 SW 07.5 309 o .f) V.0 :0 0.0 FI L LER M ~.2 Sw = 07.5 _~_~.!t~.O ____ . __ Q~ ________ 0.0 F Lt..~ ____ . ____________ 1i._~2_~~......5.

0.0 0.0 :0 ::.

311 0.0 F lLLE k M 2.2 sw 67.5 :: 317 0.0 0.0 0.0 F I LLE R M = 2.2 SW 67.5 0.0 O.Q :: ::.

313 0.0 FILLfR M 2.2 SW 67 .. 5 (J.O 314 0.0 0.0 FlLLE K M 2.2 SW 67.5 = = 315 0.0 (J •. u 0.0 FI LLE:/'\ ::: M 2.2 Sw 67.5 = 401 0.03539b 0.00b037 ALPliA::.O ::.

-0. uO,+j"o ~ETA WL 75/'1 2..2 SW :: ... o.h2.

100 Added Oct 1981 TABLE J-VII. Concluded 402 -u.000724 -O.OOOlo't 0.00001:>9 BETA ALPhA wL 15M'" 2.2 sw ;. 67.~ 403 -().001j37 -0.000317 0.000;'10 DELTA H PRIME WL 15M :; 2.2 Slol :; 07.5

404 -O.OOUlle -O.OOOO~O 0.000013 DelTA SPOILER WL 751'1 = 2.2 SW = 67.5

405 O.003~0~ 0.0009~O -0.001193 DELTA RUD U~ wL 751'1 :; l.2 SW :; 67.5 40b O.O(jvbb'1 O.0()00:'9 -0. vOu174 DELTA RUO LOW WL15M:; 2.2 sw :; 01.5 401 -0.002'198 -0.001323 0.001025 RULl VELOC P ~M:; 2.2 Sw =Jll..2 --408-o~o29Til· -O~0064-7·1--=o--=-o04(iT5·YAwvEi.oc R----WL 75M = 2.2 SW = 67.5 409 V .0 v. 0 0·. v F ILLE R M = 2.2 S W '" 0 1 • 5 4100.U 0.0 0.(; FILLER M=l.2SW.=61.5 411 0.0 0.0 0.0 FILLER M = 2.2 SW = 67.5 412 0.0 o.v 0.0 FILLEk M '" 2.2 SW :; 67.5 _. __ . __ !+.J~~ __ JJ_~.~. _______ .9~Q __ . ___ 0 __ 0 ______ El~.bJ;JL _____ . ___ ._._n ___ .•.. _t1. =._2~._~W = 67.5 414 0.0 0.0 0.0 FILLER M :: 2.2 Sw :; 67.5 415 0.0 0.0 0.0 FILLER M = 2.2 Sill = 67.5 501 -O.CO}'102 :-0.001250 0.0 ALprA.:O JVERn~A\') .. 1'1 =4:.2. Sw£=:.07 .. 5

502 0.001489 o.oooz:n ·o~o ALPhA (VERTICAL) M = 2.2 SW = 67.5

503 0.000056 0.0 0.0 ROLL VEL P (LATERAL)r\ = 2.2 SW = 67.5

504 -0.006006 -o.00ll04 -0.00057 .. Bt:TA (LATERAL) M = 2.2 sw ;. 67.5 ---.-- --~o5---ci-:0 0.0 0.0 FILLER M ;. 2.2 SW :: 67.5 506 0.0 0.0 0.0 FILLEk M = 2.2 SW :: 67.5 507 __ 0.!9 .. .. ___ ._Q~9_. ___ . _ __ 9_~ ____ . __ .H_LJ,..~!L M._ ~_ 2 .•. 2.. .. SW ::.(>].~ 508 0.0 0.0 0.0 FILLER M :; 2.2 SW :: 67.5 509 0 • 0 0 • a a • v F ILL E R M : Z • 2 S W ::: b 7 • 5 510 0.0 0.0 0.0 FILLER M ::: 2.2 Sw ::: 67.5 511 0.0 0.0 0.0 FILLER M ::: 2.2 SW = 67.5 512 0 • 0 a • 0 0 • 0 F I L LE R M ::: 2.2 S W = 67. 5 513 0 .0 0.0 . O. 0 F lL lE R . _ _ __ 1L~_Z .. LS W---=-_b_1....2.

... ··--514-0:O-----0~O----- 0.0 FIl.LER----· -- M = 2.2 SW ;. 61.5

515 0.0 0.0 O.V FIllER M = 2.2 SW = 67.5

:; 601 0.008028 0.001533 0.0 AlPHA=O (VERTICAL» M 2.2 SW ::: 67.5 602 0.000142 0.000006 0.0 ALPtiA (YER TICA l) M 2.2 sw = ·67.5 = b03 -0.000104 -0.000014 0.0 BETA ALPHA=O C/O LATM = 2.2 SW = 07.5 004 -0.000025 0.0 0.0 BETA ALPHA CLAT)M :::

c/o 2.2 sw = 67.5

005 0.000024 0.0 H 0.0 --.Qf.LTA PB IME

ILAIl M Z.Z SW = ~l.S

=

b06 0.000041 0.0 6.0 DEL TA RlD (LAT, LOW H 2.2 SW = 67.5

=

:; b07 0.000179 0.000022 0.000010 ROLL VELaC ITY PLAT M 2.2 SW

= 67.5

008 -o.0014t22 -0.000144 -0.000124 BETA H SW 'LATERAL) 2.2 ::: 61.5 = 009 0.0 0.0 ::: 0.0 FILLER M SW ::: 67.5 2.2 610 0.0 0.0 0.0 FILLER M 2.2 SW ::: 07.5

=

011 0.0 0.0 0;.0 ::: FIllER M 2.2 sw = 67.5.

012 0.0 0.0 0.0 FILLER M 2.2 SW ::: 01.5

=

::: b13 0.0 0.0 0.0 FILLE R M 2.2 SW = 07.5 014 0.0 0.0 ::: 0.0 FILLER M sw :; 61.5 2.Z 0~5 0.0 0.0 0.0 FILLE R M 2.2 SW == 07.5 = Added Oct 1981 1. Report No. 2. Government Accession No.

3. Recipient's Ciltalog No.

NASA CR-170410 4. Title and Subtitle 5. Report Date January 1984 AIRLOADS RES~ARCH STUDY 6. Performing Organization Code VOX,UME II: AIRLOAD COEFFICIENTS OERIVED FROM WINO TUNNEL DATA 7. Author(s) 8.

Performing Organization Report No.

D. Olsen, Jr., M. O. Bartlett, T. F. Feltz, A. NA-76-563 D. B. Smith, and P. F. Wildermuth 10. Work Unit No.

9. P!)rforrning Organization Name and Address Rockwell International 11. Contract or Grant No.

LoS Angeles Division NAS4-2769 Los Angeles, California 90009 13. Type of Report and Period Covered Sponsoring Agency Name and Address 12.

Contractor Re~ort - Final National Aeronautics and Space Administration 14. Sponsoring Agency Code washington D.C. 20546 RTOP 505-33-54 15. Supplementary Notes Facility, NASA Technical Monitor: Robert L. Sims, Ames Research Center, Dryden Flight Research Edwards, CA. 93523. "Volume I: Flight Test Loads Acquisition" is published as NASA CR-170409.

16. Abstract This report describes the development of B-1 aircraft rigid wind tunnel data for use in subsequent tasks of the Airloads Research Study. The basic intent of the overall program is to utilize data acquired during B-1 aircraft tests, analyze these data beyond the scope of Air Force requirements, and pre- pare research reports that will add to the technology base for future large flexible aircraft.

from the Rockwell International external structural loads data bank Data were used to generate coefficients of rigid airload shear, bending moment, and torsion at specific component reference stations for both symmetric a~d asym- Component stations include the movable wing, horizontal and metric loadings.

vertical stabilizers, and forward and aft fuselages. The coefficient data cover a Mach number range from 0.7 to 2.2 for a wing sweep position of 67.5 • 17. Key Words (Suggested by Author(s)) 18. Distribution Statement Unclassified-Unlimited B-1 airplane Wind tUnnel airloads STAR category 05 Security Classif. (of this page) Price" Sec\Jrity Classif. (of this report) 20. 21. No. of Pages 22.

19.

Unclassified Unclassified 107 A06 *For sale by the National Technical Information Service, Springfield, Virginia 22161 GPO 788-389/137

End of Document