NASA TECHNICAL NOTE NASA TN 0-5971
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LONGITUDINAL AERODYNAMIC CHARACTERISTICS OF A TWIN-TURBOFAN SUBSONIC TRANSPORT WITH NACELLES MOUNTED UNDER THE WINGS
by Francis J. Capone
Langley Research Center
Hampton, Va. 23365
NATIONALAERONAUTICSAND SPACEADMINISTRATION• WASHINGTON,D. C. ° OCTOBER1970
1, Report No, 2, Government Accession No. 3, Recipient's Catalog No. NASA TN D-5971
4, Title and Subtitle 5. Report Date LONGITUDINAL AERODYNAMIC CHARACTERISTICS OF A October 1970
TWIN-TURBOFAN SUBSONIC TRANSPORT WITH NACELLES 6. Performing Organization Code MOUNTED UNDER THE WINGS
7. Author(s) 8. Performing Organization Report No.
Francis J. Capone L-7141
10. Work Unit No. 9. Performing Organization Name and Address 737-01-10-03
NASA Langley Research Center 11. Contract or Grant No Hampton, Va. 23365 13, Type of Report and Period Covered
12. Sponsoring Agency Name and Address Technical Note National Aeronautics and Space Administration 14. Sponsoring Agency Code Washington, D.C. 20546
15. Supplementary Notes
16. Abstract
An investigation has been conducted in the Langley 16-foot transonic tunnel to deter- mine the longitudinal aerodynamic characteristics of a 0.062-scale, twin-turbofan subsonic transport at Mach numbers from 0.55 to 0.85 and angles of attack from about -2 ° to 6 °. The Reynolds number based on wing mean aerodynamic chord varied from 2.25 × 106 to 2.70 × 106. The effects of model-component buildup, horizontal-tail effectiveness, boundary- layer transition, and wing and nacelle modifications were measured. The model was mounted by using a sting-strut arrangement with the strut entering the model through the underside of the fuselage approximately 65 percent of the fuselage length rearward of the model nose. Strut-interference effects were measured and applied as a correction to the data.
t7. Key Words (Suggested by Author(s)) 18. Distribution Statement
Subsonic _erodynamics Unclassified - Unlimited Subsonic transport Strut interference
19. Security Classif. (of this report) 20. Security Classif. (of this pagel 21. No. of Pages 22. Price"
Unclassified Unclassified 93 $3.00
"For sale by the Clearinghouse for Federal Scientific and Technical Information
Springfield, Virginia 22151
LONGITUDINAL AERODYNAMIC CHARACTERISTICS OF
A TWIN-TURBOFAN SUBSONIC TRANSPORT WITH
NACELLES MOUNTED UNDER THE WINGS
By Francis J. Capone Langley Research Center
SUMMARY
An investigation has been conducted in the Langley 16-foot transonic tunnel to deter- mine the longitudinal aerodynamic characteristics of a 0.062-scale, twin-turbofan sub- sonic transport at Mach numbers from 0.55 to 0.85 and angles of attack from about -2 ° to 6 ° . The engine nacelles were mounted under the wings. The Reynolds number based on wing mean aerodynamic chord varied from 2.25 x 106 to 2.70 x 106 . The effects of model-component buildup, horizontal-tail effectiveness, boundary-layer transition, and wing and nacelle modifications were measured. The model was mounted by using a sting- strut arrangement with the strut entering the model through the underside of the fuselage approximately 65 percent of the fuselage length rearward of the model nose. Strut- interference effects were measured and applied as a correction to the data.
For the small range of tail deflection (-0.5 ° to 0.5o), there was little or no effect of horizontal-tail deflection on lift-curve slope, model stability, drag coefficient, or maxi- mum lift-drag ratio. The model with free boundary-layer transition had more lift at high angles of attack and less stability; and for tail deflections of 0 ° and 0.5 °, the lift coeffi- cient at which pitchup instabilities occurred was higher than that for the model with fixed transition. The configuration with a modified wing (reduced wing thickness ratio) and longer nacelles had greater lift at the same angle of attack and a higher lift coefficient at which pitchup instabilities occurred than did the basic configuration. The drag-rise Mach number for the modified configuration was increased by 0.02, and the modified config- uration had much less drag due to lift at the higher Mach numbers than that of the basic configuration.
INTRODUCTION
The National Aeronautics and Space Administration has conducted a wind-tunnel investigation to obtain data for a correlation between wind-tunnel and flight-test results for a twin-turbofan, short-haul, subsonic transport with engines mounted under the wings. This airplane is capable of carrying about 100 passengers. This report presents only the results of the wind-tunnel investigation. The wind-tunnel investigation was conductedwith a 0.062-scale model in the Langley 16-foot transonic tunnel at Mach numbersfrom 0.55to 0.85and angles of attack from -2o to about6°. The Reynoldsnumber basedon wing meanaerodynamicchord varied from 2.25× 106to 2.70 x 106. The effects of model-componentbuildup, horizontal-tail effec- tiveness, boundary-layer transition, andwing andnacelle modifications were measured. Subsonictransports are frequently designedwith fuselage aftersections that are nonsymmetrical with a large amountof upsweepon the bottom of the afterbody. The pressure drag on this section of the fuselage can be a large part of the total fuselage drag. A recent investigation as reported in reference 1 was concernedwith evaluatingthe sting- support interference effects of a conventionalstraight sting that enteredthrough the rear of three subsonictransport models. However,the straight sting required a rather large cutoUton the afterbody of the fuselage. The model of the present investigation was mountedin the windtunnel by using a sting-strut support arrangementwith the strut entering the modelthrough the underside of the fuselage approximately 65 percent of the fuselage,length rearward of the model nosewhich minimized alterations to the fuselage. Strut-support interference effects were determined andapplied as a correction to the measuredaerodynamiccharacteristics.
SYMBOLS
Model forces and momentsare referred to a stability axis system with the model moment reference center located 82.80centimeters rearward of the model nose corre- spondingto 22.4percent of the wing meanaerodynamicchord which is approximately at the nominal center-of-gravity position of the airplane. Dimensions are given in the International Systemof Units (SI).
A aspect ratio
local wing chord
wing or tail mean aerodynamicchord (Wing _ = 21.17 centimeters)
CA,i nacelle internal axial-force coefficient
CD drag coefficient, Drag qS
CD,p computed profile drag coefficient
CD,min minimum drag coefficient
2 CD,trim trim drag coefficient
CL lift coefficient, Lift qS
CL,M lift coefficient at CD,min
CL,trim trim lift coefficient (lift coefficient at Cm = 0)
CL_ lift-curve slope per degree
C m pitching-moment coefficient, Pitching moment qS_
0Cm Cmc L static-longitudinal-stability parameter, OCL
Cm, o pitching-moment coefficient at zero lift
average drag-coefficient increment due to strut interference, (CD)with strut- (CD)without strut
Z_CD,HV drag-coefficient increment due to adding horizontal and vertical tails
drag-coefficient increment due to adding nacelle 1 and pylon
average lift-coefficient increment due to strut interference, (CL)with strut-(CL)without strut
ACm,av average pitching-moment-coefficient increment due to strut interference, (Cm)with strut-(Cm)withoat strut
dC D k M drag-due-to-lift factor, d(CL- CL,M) 2 L/D lift-drag ratio
(L/D)max maximum lift-drag ratio
(L/D)tri m lift-drag ratio at trim conditions
M free-stream Much number q free-stream dynamicpressure
R Reynoldsnumber per meter
wing reference area (3674.30centimeters2)
Tt stagnation temperature
X,Z wing coordinates
Ot w wing angle of attack (1 ° with respect to body center line)
6h incidence angle of horizontal tail, positive when trailing edge is down
Subscripts: l lower u upper
Abbreviations:
LER leading-edge radius
WL water line
Model- component designations:
B fuselage plus wing-root-flap actuator fairing
H horizontal tail
N1 basic nacelle and pylon
N2 basic nacelle and pylon with rear-end extension
T wing trailing-edge-flap actuator fairings (two each located outboard on wing)
V vertical tail
4 W1 basic wing
W2 basic wing plus leading- andtrailing-edge chord extensions
W3 basic wing plus trailing-edge chord extension
APPARATUS
Model
The complete 0.062-scale basic model is shown in the sketch and photographs of figures 1 and 2, respectively. The model represented a twin-turbofan, short-haul, sub- sonic transport weighing about 45 000 kilograms that was capable of carrying about 100 passengers at a cruise Mach number between 0.78 and 0.80. The design lift coeffi- cient is 0.30. Details of the various model components are presented in figure 3.
Fuselage.- Fuselage geometry and cross sections are shown in figures 3(a) and 3(b), respectively. The fuselage was 171.19 centimeters long and had a fineness ratio of 6.9 based on the maximum body depth. The wing root fairing was located between fuse- lage stations 54.33 and 109.76. A fairing on the fuselage used to house a wing trailing- edge flap track and actuator mechanism extended from station 85.04 to 105.03. (See fig. 3(b).) The wing trailing-edge flaps were not simulated during this investigation.
Basic wing.- The planform geometry of the basic wing (Wl) is shown in figure 3(c). The basic wing had an aspect ratio of 8.41, a span of 175.59 centimeters, an incidence angle of 1 °, and a dihedral angle of 6 °. Both the leading and trailing edges had discon- tinuous sweep. Airfoil ordinates for the basic wing are presented in table I(a). The model with the basic wing is shown in the photographs of figure 2.
Two modifications were made to the basic wing as shown in figure 3(d). The modi- fication for wing 2 (W2_ consisted of a 1-percent-chord leading-edge extension and a 15-percent-chord trailing-edge extension, both outboard of span station 32.19 which was the spanwise location of the break in the leading and trailing edges of the basic wing. The trailing edge also extended inboard to the nacelle pylon. Photographs of the model with this wing are presented in figure 4. The modification for wing 3 (W3) involved only a trailing-edge extension from 15 percent chord at span station 32.19 to 0 percent chord at span station 65.67. The model with this wing is shown in the photographs of figure 5. Airfoil ordinates for wings 2 and 3 are presented in table I(b). These ordinates were non- dimensionalized with respect to the local chords of wing 1. Therefore, these modifica- tions reduce the wing thickness ratio when based on the chord of the modified wing. For example, at span station 32.19, the maximum wing thickness ratio for wing 1 (based on the chord of wing 1) is 0.108, and for wing 3 the maximum thickness ratio is 0.095 (based on the chord of wing 3). It should be notedthat wings 2 and3 were tested with a different nacelle from that on wing 1 andwithout the flap-track fairings on the wings. The flap- track fairing at the wing root, however,was present. Only the complete modelusing wing 1 includedfairings on the wing for the wing flap tracks andactuators. A sketch of the fairings is presented in figure 3(e). These fairings were located at spanstations 40.64and 56.82 and are shown in the photographs of figure 2.
Nacelles.- Sketches of the two nacelles tested are presented in figure 3(f).
Nacelle 1 (N1) was 34.54 centimeters long and was tested only with wing 1. Nacelle 2 (N2) was similar to nacelle 1 except that the rear portion was extended 7.08 centimeters resulting in a total length of 41.62 centimeters. This nacelle was tested only with wings 2 and 3. Both nacelle inlets had the same geometry and were located at the same body station.
Horizontal and vertical tails.- Figures 3(g) and 3(h) show the planform geometry of the horizontal and vertical tails, respectively. Airfoil ordinates are presented in tables H and III. The horizontal tail was all-movable with the hinge axis located at fuse- lage station 160.93.
Model Support System
The present investigation utilized a sting-strut mount in order to minimize the alterations made to the fuselage for a support system. A sketch showing the various support systems is presented in figure 6. For determining the aerodynamic characteris- tics of the model, the model was supported with the strut entering through the underside of the fuselage at a location approximately 65 percent of the fuselage length to the rear of the nose as shown by the sting-strut arrangement of figure 6(a). This mounting sys- tem is also shown in the photographs of figures 2, 4, and 5. This type of strut allowed for the minimum amount of cutout to the model (as compared with the large amount of cutout to the models of ref. 1) since the strut chord length at the body juncture was about 25.4 centimeters with a maximum thickness of about 2.54 centimeters.
In order to assess the magnitude of the strut interference, two additional support systems were used as shown in figures 6(b), 6(c), and the photographs of figure 7. Fig- ures 6(b) and 7(a) show the model with the strut entering through the top of the model. A dummy sting strut was attached to the live sting strut through a blade downstream of the model base. The dummy strut entered through the bottom of the model (at the same location as the live strut). A positioning pin that was part of the dummy strut fit loosely into the balance support block and was the only point of contact inside the model. The loose fit of the pin allowed model deflection with the same aeroelastic support stiffness as existed with only the live strut present. The support system of figures 6(c) and 7(b) showsthe model with only the live strut entering through the top of the model (dummy sting strut removed). The gapsbetweenthe struts andmodel were sealed with synthetic spongerubber. Pressure inside the model was continuouslymonitored in order to detect andwarn of possible leakagethrough the seal if it occurred. Calibrations of normal force, axial force, andpitching moment with the model assembledshowedno restraint due to this method of sealing. It should be notedthat the vertical tail could not be attachedwhile using the top mount system. Becauseof the mounting arrangements, the model was tested aboveandbelow the wind-tunnel center line (fig. 6). Therefore, it was necessary to test the wing-body com- bination upright and inverted in both wind-tunnel positions in order to determine the mag- nitude of the wind-tunnel flow angularity. This was accomplishedby using the sting- strutmdummy-strut combination with the live strut coming into the model from either the top or bottom as shownin the following table:
Model position to tunnel Live strut Dummy strut Model attitude center line
Below Top Bottom Upright Below Bottom Top Inverted Above Top Bottom Inverted Above Bottom Top Upright
Wind Tunnel and Instrumentation This investigation was conductedin the Langley 16-foot transonic wind tunnel which is a single-return, atmospheric wind tunnel with a slotted octagonaltest section andcon- tinuous air exchange. For models mountedalongthe tunnel center line, the model- support angle-of-attack mechanismpivots the sting support in such a manner that the model is close to the center line. However, for the present investigation with its sting- strut arrangement, that puts the model either aboveor below the tunnel center line, there is some translation of the modelalong with the rotation. The center of model rotation is indicated in figure 6. Aerodynamic forces were measured with an internally located, six-component strain-gage balance. Angle of attack was determined with a pendulum-type strain-gage inclinometer located inside the model nose. For the determination of nacelle internal axial force, stagnation pressures at the nacelle-duct exit were measured on a pressure- scanningunit; whereas static pressures at the nacelle-duct exit were measured with indi- vidual pressure transducers. TESTS
This investigation was conductedat Mach numbersfrom 0.55to 0.85 and at wing angles of attack from -2° to about6°. The Reynoldsnumber basedon the mean aero- dynamic chord varied from 2.25 × 106to 2.70 x 106. All model configurations except as notedwere tested with boundary-layer transition strips consisting of No. 120silicon carbide grit particles sparsely distributed in a thin film of lacquer that was 0.25 centi- meter wide. These strips were located on both the upper andlower surfaces of the wings andtails at 10 percent of the local streamwise chord, onthe nacelles (outside and inside) at 0.76 centimeter from the nacelle leading edge,and on the fuselage nose at 2.54 centi- meters from the tip of the nose. The grit size andthe location of the strips were deter- mined according to the recommendationsof reference 2. The aerodynamiccharacteristics of the various model configurations were deter- minedwith the model mountedin the wind tunnel as shownin figures 2, 4, 5, and6(a), that is, with the strut entering the modelfrom the bottom. The effects of model- componentbuildup, horizontal-tail deflections (5h = 0° and _0.5°), boundary-layer transi- tion not artificially fixed, andtwo wing andnacelle modifications were studied. Sting-strut interference effects were measuredby testing configuration BWlH (5h = -0.5 °) which was supported by the three methods shown in figure 6. Wind-tunnel flow angularity was determined by conducting tests of configuration BWl, upright and inverted, both above and below the wind-tunnel center line. Both configurations BWlH and BW 1 were tested with transition fixed. The methods of support for this portion of the tests were summarized in a previous section entitled "Apparatus." Nacelle internal axial force was determined from measurements of both static and stagnation pressure at the exit of one nacelle by means of a pressure survey rake that was rigidly attached to the nacelle.
CORRECTIONS AND ACCURACY
General Corrections
The wind-tunnel flow angularity as determined by tests of the model upright and inverted, both above and below the wind-tunnel center line, was found to be 0 °. The mea- sured balance axial force was corrected for nacelle internal axial force shown in fig- ure 8. The effects of nacelle incidence and the variation of CA, i with angle of attack were accounted for in applying the correction to the balance axial force.
No corrections have been made for roughness drag due to the grit applied for the boundary-layer transition strips. These corrections are considered unnecessary at sub- sonic speeds since the general guideline for the application of transition strips (ref. 3) and a grit height based on a transitional Reynolds number of 600 were used. (See also ref. 2.) Corrections to the lift data from either solid-blockage interference or tunnel- boundary interference effects are not considered necessary. Theoretical calculations presented in reference 4 for a model with approximately the same wing span and cross- sectional area have shown that the tunnel-wall lift-interference correction reduced the angle of attack by 0.02C L. For the model of reference 4, the reduced angle of attack at C L = 0.5 reduced the drag coefficient by 0.0001. Since there is no Mach number gradi- ent in the Langley 16-foot transonic tunnel, no corrections for buoyancy are made.
Since the gap between the strut and the model was sealed with the synthetic sponge rubber, no correctiona are necessary for either the base or balance cavity. An advan- tage of this type of mounting over that of reference I can be seen in the application of corrections to the measured aerodynamic characteristics due to the effects of the sting cavity. It was necessary in the investigation of reference 1 to measure the longitudinal variation of cavity pressure along the sting and fuselage-sting cavity. In addition to cor- recting the axial force (to the condition of free-stream static pressure acting across the sting cavity), an adjustment is necessary to both normal force and pitching moment. This was done by integrating the pressures along the sting longitudinally and obtaining incre- mental corrections.
Sting-Strut Interference Correction
The technique used to determine the sting-strut interference effects was similar to that described in reference 1. Force and moment measurements made with the model supported with the strut through the bottom of the model (fig. 6(a)) will, of course, contain an interference term of the bottom strut on the model that must be subtracted from the measurements made. This interference term can be evaluated from measurements made when the model is supported as shown in figures 6(b) and 6(c).
When the model is supported with the strut through the top and the dummy strut in place from the bottom (fig. 6(b)), the measured forces contain interference terms caused by both the top and bottom struts. The interference term due to only the top strut is con- tained in the measured data when the model is supported as shown in figure 6(c). There- fore, subtracting the coefficient data measured when the model is supported as shown in figure 6(c) from that of figure 6(b) will result in the desired strut-interference term.
The variation of strut-interference terms for lift, drag, and pitching moment with wing angle of attack for the Mach numbers investigated is presented in figure 9. Shown are average faired values for each of these components. Corrections to the measured aerodynamic data were made automatically when processing the data by computer by inputing a table of the strut-interference terms as a function of the wing angle of attack at 0.25 ° increments and linearly interpolating between input points.
9 Accuracy The accuracy of data presented herein prior to making corrections for strut inter- ference, basedprimarily on expectedinstrumentation accuracies, is presented as follows: M ...... i0.007 aw, deg ...... +0.10 C L ...... +0.004 C D ...... ±0.0005 C m ...... +0.003
5h, deg ...... +0.03
PRESENTATION OF RESULTS
The basic longitudinal aerodynamic force and moment coefficients are presented in figures 10 to 14 as follows:
Configuration 5h, deg Transition Figure
BW 1 and BW1HV -0.5 Fixed 10
BW1HVN1T -.5 Fixed and free 11 BWlHVN1T 0 Fixed and free 12 BWlHVN1T .5 Fixed and free 13 BWlHVN 1 , BW2HVN2, -.5 Fixed 14 and BW3HVN 2
Various summary plots of the longitudinal aerodynamic characteristics are pre- sented in figures 15 to 19 as follows: Figure Effect of model-component buildup ...... 15 Computed profile drag coefficients ...... 16 Effect of horizontal-tail deflection ...... 17 Effect of fixing boundary-layer transition ...... : ...... 18 Effect of wing and nacelle modifications ...... 19
DISCUSSION
Effect of Model-Component Buildup
The effects of model-component buildup can be seen by comparing the basic data presented in figures 10, 11, and 14 and the summary data of figure 15 for configurations
BWl, BWlHV , BWlHVN1, and BWlHVN1T with 5h = -0.5 ° and the transition fixed. The
10 lift curves for these configurations are linear over the angle-of-attack range investigated at Mach numbers of 0.55 and 0.625. At M = 0.725 or greater, nonlinearities in the lift curves occurred at lower liftcoefficients as Mach number was increased.
The pitching-moment curves exhibited nonlinearities at approximately the same value of C L as the liftcurves with the model either becoming neutrally stable (except
BWl) or having a pitchup instability.
Lift-curve slope CLo l increased by about 10 percent up to M = 0.825 with the addition of the tail surfaces (fig. 15(a)). The nacelles caused a further small increase in CL_ up to M = 0.80, whereas there was no effect on CLo l of adding the flap-track fairings.
The nacelles and flap-track fairings together (compare configurations BWIHV and
BWIHVNIT ) are seen to cause a reduction in both the stability Cmc L and in Cm, o (fig. 15(a)). This reduction in Cm, o is 0.026 at M = 0.55 and 0.039 at M = 0.80. This reduction in Cm, o results in CL,trim being reduced from 0.355 to 0.320 at M = 0.55, whereas CL,trim is reduced from 0.5 to 0.3 at M = 0.80. (See figs. i0 and 11.)
The effects on minimum drag and drag at lifting conditions caused by the various model components are shown in figure 15(b). Computed profile drag coefficients and measured incremental drag coefficients for the tail surfaces and the nacelles are pre- sented in figure 16. Skin-friction drag coefficients were computed by the methods of references 5 and 6. Wetted areas, reference lengths, average Reynolds number per meter, wind-tunnel stagnation temperature, and computed CD, p are presented in table IV. Form factors given in chapter XXIV of reference 7 were used to obtain the pro- file drag coefficients from the skin-friction drag coefficients. The incremental drag due to a particular model component is nearly constant with Mach number up to M = 0.775 at each of the liftcoefficients presented (fig. 15(b)). Addition of the tails caused ml
increase in CD,mi n from 0.0049 at M = 0.55 to 0.0060 at M = 0.85 (fig. 16). This drag increment was 0.0008 to 0.0020 higher than the computed profile drag coefficient, probably due to some tail drag due to liftand interference of the tails on the fuselage afterbody.
The same observations can be made for the addition of the nacelles where an inter- ference drag of about 0.0006 is indicated up to about M = 0.775 at minimum drag condi- tions (fig. 16). Both of the incremental drag coefficients caused from adding either the tails or nacelles indicate drag-rise Mach numbers of about 0.775. The drag-rise Mach number (where dCD,min/dM = 0.i) is between 0.78 and 0.80 for the complete configura- tion BWlHVNIT (fig. 15(b)).
II The maximum lift-drag ratio for configuration BW 1 was 20.2. (See fig. 15(c).) A loss in L/D of about 4 occurred due to the addition of the remainder of the model com- ponents and resulted in a value of (L/D)max of 16.2 for configuration BWlHVN1T.
Effect of Horizontal-Tail Deflection
The effect of horizontal-tail deflection for the complete configuration BWlHVN1T with transition fixed or free can be seen by comparing the basic data of figures 11, 12, and 13 and the summary data of figure 17 for the condition of transition fixed.
Horizontal-tail deflection had little or no effect on lift-curve slope and only a small effect on the model stability (fig. 17(a)). The effects due to this small range of tail deflection on CD,mi n or CD at lifting conditions (fig. 17(b)) and on (L/D)max (fig. 17(c)) were small as expected. Trimmed drag polars are presented in figure 17(d) where the symbols represent the trim points obtained from the pitching-moment data of figures 11, 12, and 13 for the transition-fixed conditions.
Effect of Boundary-Layer Transition
The basic aerodynamic data for configuration BWlHVN1T for three horizontal-tail deflections with boundary-layer transition fixed and free are presented in figures 11, 12, and 13, and summary data are presented in figure 18. Generally, fixing transition had little or no effect on CLa up to M = 0.825. Lift-curve slopes shown in figure 18(a) for 5h = -0.5 ° are a typical example. The model with free transition had more lift at high angles of attack at Mach numbers greater than 0.75 for the three tail deflections.
For the model with 5h = -0.5 ° (fig. 11), the pitching-moment curves were more linear over a greater range of lift coefficient with transition fixed. Pitchup occurs at approximately the same lift coefficient. However, for the model with 5h = 0o or 0.5 ° (figs. 12 and 13, respectively), the pitching-moment curves were linear over the same range of lift coefficient with transition both free and fixed, and the lift coefficient at which the pitchup instability occurred was substantially higher.
Shown in figure 18(a) is Cmc L for the three tail deflections with transition free and also Cmc L for 5h = -0.5 ° with transition both fixed and free. The variation Cmc L with transition fixed and free at the other two tail deflections (0 ° and 0.5 °) is similar to that shown in figure 17(a). Figure 18(b) presents Cm, o data with transition fixed and free for the three tail deflections. The model with transition free exhibited lower stability and higher values of pitching-moment coefficient at zero lift. A compari- son of Cmc L curves with transition fixed and free (figs. 18(a) and 18(b)) shows a rear- ward shift in the center of pressure due to fixing transition from 5 to 10 percent of the mean aerodynamic chord at Mach numbers up to 0.75.
12 The effects on drag and lift-drag ratio of fixing transition are compared only for 5h = -0.5 ° in figures 18(c) and 18(d). The results for the other two tail settings are similar. As expected, the free-transition condition exhibited lower drag and higher (L/D)max than the fixed-transition condition.
Effect of Wing Modification
A comparison of the aerodynamic effects due to the two wing modifications for the complete model without the flap-track fairings, with transition fixed and 5h = -0.5 °, is shown by the basic data of figure 14 and the summary data of figure 19. These modifica- tions, made to the basic wing (Wl) outboard of the nacelle (span station 30.07), were intended to increase aerodynamic performance between the cruise Mach numbers of 0.78 and 0.80 and possibly to increase the cruise Mach number. Wing 2 (W2) had both a 1-percent-chord leading-edge extension and a 15-percent-chord trailing-edge extension which resulted in the leading-edge sweep being increased from 27.58 ° to 27.68 °. Wing 3 (W3) had only a 15-percent-chord trailing-edge extension at span station 32.19 which tapered to 0 percent chord at span station 65.67. (See figs. 3(c), 3(d), 4, and 5.) Both wings 2 and 3 were tested with the longer nacelle. (See fig. 3(f).)
Configuration BW2HVN 2 had a higher lift-curve slope CL_ over the entire Mach number range than that of BWlHVN1, whereas CL_ for configuration BW3HVN 2 was higher only at Mach numbers greater than 0.775 (fig. 19(a)).
Except at the lower Mach numbers and at M = 0.85, the three wings had about the same stability (fig. 19(a)). However, the modifications to the basic wing increased the lift coefficient at which pitchup instabilities occurred (fig. 14). Configuration BW2HVN 2 showed no pitchup tendencies until M = 0.775.
The drag characteristics of figure 19(b) show configuration BW2HVN 2 to have lower drag as Mach number and lift coefficient are increased. At C L = 0.3 and M = 0.8 the drag coefficient of this configuration is 0.0058 less than that of configuration BWlHVN 1. The drag-rise Mach number for configuration BW2HVN 2 is about 0.82 as compared with a Mach number range from 0.78 to 0.80 for configuration BWlHVN 1. It should be noted that the aerodynamic coefficients presented in figures 14 and 19 were computed based upon the reference wing area of configuration BWlHVN 1. Using the actual wing area, for example, of configuration BW2HVN 2 would decrease the drag coefficients by some 7 percent.
From the lift-drag polars of figure 14 it is obvious that significant changes in drag- due-to-lift factor, especially at high Mach numbers, were a result of the wing modifica- tions. Therefore, the drag-due-to-lift factor k M for each configuration has been deter- mined by fitting the equation
13 C D = CD,mi n + kM(C L - CL,m)2
to the lift-drag polars. (See refs. 8 and 9.) The resulting drag-due-to-lift factors are
presented in figure 19(c) in product form kMA. In this form, kMA is independent of the wing reference area used in calculating the two parameters. This equation was fitted
over a range of lift coefficients from about -0.1 to the lift coefficient just below C L for
(L/D)ma x. Also presented in figure 19(c) are the conditions for zero suction A/CLa , and for 100 percent suction 1/_.
The abrupt rise in the drag-due-to-lift factor for configuration BW2HVN 2 occurs at M = 0.80 which is 0.05 higher than that for the other two configurations. This prob- ably results from the lower thickness ratio of the modified wing and the higher critical Mach number of the airfoil.
Lift-drag-ratio characteristics are shown in figure 19(d). Up to M = 0.725, wing modifications have little or no effect on (L/D)max. However, at higher Mach numbers, the modified wings display higher values of (L/D)max than those of the unmodified
wing. For example, at M = 0.80, (L/D)max for configuration BW2HVN 2 was 3.4 higher than that for BWlHVN 1.
SUMMARY OF RESULTS
An investigation has been conducted in the Langley 16-foot transonic tunnel to
determine the longitudinal aerodynamic characteristics of a 0.062-scale, twin-turbofan subsonic transport at Mach numbers from 0.55 to 0.85 and angles of attack from about -2 ° to 6 ° . The Reynolds number based on wing mean aerodynamic chord varied from 2.25 x 106 to 2.70 × 106 . The effects of model-colnponent L'uildup, horizontal-tail
effectiveness, boundary-layer transition, and wing and nacelle modifications were mea- sured. The model was mounted by using a sting-strut arrangement with the strut entering the model through the underside of the fuselage approximately 65 percent of the fuselage length rearward of the model nose. Strut-interference effects were measured and applied as a correction to the data. The investigation indicated the following results:
1. For the small range of tail deflection (-0.5 ° to 0.5o), there was little or no effect of horizontal-tail deflection on lift-curve slope, model stability, drag coefficient, or maximum lift-drag ratio.
2. The model with free boundary-layer transition had more lift at high angles of attack and less stability; and for tail deflections of 0 ° and 0.5 °, the lift coefficient at which pitchup instability occurred was higher than that for the model with fixed transition. The model with free transition had lower stability over the Mach number range tested.
14 3. The configuration with a modified wing (reduced wing thickness ratio) and longer nacelles had greater lift at the same angle of attack and a higher lift coefficient at which pitchup instabilities occurred than that of the basic configuration. The drag-rise Mach number for the modified configuration was 0.02 higher than that of the basic configura- tion, and the modified configuration had much less drag due to lift at the higher Mach numbers.
Langley Research Center, National Aeronautics and Space Administration, Hampton, Va., June 19, 1970.
15 RE FERENCES
1. Loving, Donald L.; and Luoma, Arvo A.: Sting-Support Interference on Longitudinal Aerodynamic Characteristics of Cargo-Type Airplane Models at Mach 0.70 to 0.84. NASA TN D-4021, 1967.
2. Braslow, Albert L.; and Knox, Eugene C.: Simplified Method for Determination of Critical Height of Distributed Roughness Particles for Boundary-Layer Transition at Math Numbers From 0 to 5. NACA TN 4363, 1958.
3. Braslow, Albert L.; Hicks, Raymond M.; and Harris, Roy V., Jr.: Use of Grit-Type Boundary-Layer-Transition Trips on Wind-Tunnel Models. NASA TN D-3579, 1966.
4. Luoma, Arvo A.; Re, Richard J.; and Loving, Donald L.: Subsonic Longitudinal Aero- dynamic Measurements on a Transport Model In Two Slotted Tunnels Differing in Size. NASATMX-1660, 1968.
5. Schoenherr, Karl E.: Resistance of Flat Surfaces Moving Through a Fluid. Trans. Soc. Naval Architects Marine Eng., 1932, pp. 279-313.
6. Sommer, Simon C.; and Short, Barbara J.: Free-Flight Measurements of Turbulent- Boundary-Layer Skin Friction in the Presence of Severe Aerodynamic Heating at Mach Numbers From 2.8 to 7.0. NACA TN 3391, 1955.
7. Schlichting, Hermann (J. Kestin, transl.): Boundary Layer Theory. Fourth ed., McGraw-Hill Book Co., Inc., c.1960.
8. Igoe, William B.; Re, Richard J.; and Cassetti, Marlowe D.: Transonic Aerodynamic Characteristics of a Wing-Body Combination Having a 52.5 ° Sweptback Wing of Aspect Ratio 3 With Conical Camber and Designed for a Mach Number of _/2. NASA TN D-817, 1961.
9. Re, Richard J.; and Stumbris, Gunars: An Investigation at Mach Numbers From 0.40 to 1.00 of a Model With a Wing Having Inboard Sections Cambered for Mach 1.2. NASA TN D-3419, 1966.
16 0
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19 TABLE I.- WING AIRFOIL ORDINATES - Concluded
(b) Wings 2 and 3
Wing 2 span stations at Wing 3 span stations at - 32.189 and outboard - x/c ×/c 32.189 48.928 Zu/C zl/c
-0.0100 0 0 7U, C ZI/C _u/C zz/c 0 0 0 0 .0040 -.0040 0 0 .0025 .0(70 -.0051 .0070 -.0051 .0025 .0105 -.0035 .0050 30 -.0066 .0100 -.0066 .0050 .0135 -.0050 .0250 .0232 -.0116 .0232 -.0116 .0100 .0190 -.0070 .0500 .0335 -.0148 .0335 -.0145 .0250 •0284 -.0100 i .1000 .0468 -.0200 .0464 -.0206 i .0500 .0385 -.0135 .1500 .0549 -.0246 •1000 .0489 -.0190 .0551 -.0248 i .2000 .0606 -.0291 .0601 -.0286 .1500 .0559 .3000 .0670 -.0365 .0670 -.0365 .2000 ,0608 .3500 .0685 -.0390 .0680 -.0389 .3000 ,0670 -.0365 .4000 .0694 -.0405 .0687 -.0407 .3500 ,0685 -.0390 .4500 .0700 -.0400 .0687 -.0407 1 .4000 ,0694 -.0405 .5000 .0710 -.0394 .0687 -.0395 .4500 0700 -.0400 .5500 .0710 -.0375 .0673 -.0375 ! .5000 0710 -.0394 .6000 .0710 -.0340 .0650 -.0340 .5500 0710 -.0375 .6500 .0700 -.0300 .0602 -.0303 .6000 0710 -.0340 .7000 .0675 -.0259 .6500 0700 -.0300 .0576 -.0260 4 .7500 .0640 -.0223 .0526 -.0220 1 .7000 0675 -.0259 .8000 .0590 -.0177 .7500 0640 -.0223 .0464 -.0175 i .8500 .0530 -.0143 .0400 -.0129 .8000 0590 -.0177 .9000 .0465 -.0102 .0330 -.0092 .8500 0530 -.0143 .9500 .0400 -.0035 .0267 -.0056 .9000 0465 -.0102 1.0000 .9329 -.0023 .0189 -.0017 ] .9500 0400 -.0035 1.0500 .0258 .0018 1.0000 .0120 .0019 .0329 -.0023 I 1.0900 .0043 .0043 i 1.0500 .0258 .0018 1.1500 .0110 .0110 1.1000 .0183 .0054 i I
1.1500 .0110 .0110 C_ Cn] . . . 22.017 18.263
Span station 32.189, c = 22.017 cm LER/c... 0.0049 0.0049 Span station 71.928, c = 13.129 cm
18 TABLE II.- HORIZONTAL-TAIL AIRFOIL ORDINATES
Horizontal-tail span stations at - x/c 0.0 7.010 15.354to 33.381
Zu/O zUo zu/c z//c Zu/C zUc 0 0 -0.0140 0 -0.0114 0 -0.0104 .0050 .0093 -.0230 .0085 -.0197 .0080 -.0184 .0075 .0118 -.0260 .0111 -.0223 .0100 -.0213 .0125 .0156 -.0306 .0143 -.0267 .0128 -.0254 .0250 .0209 -.0387 .0182 -.0347 .0164 -.0324 .0500 .0251 -.0489 .0208 -.0448 .0186 -.0422 .0750 .0270 -.0564 .0222 -.0516 .0200 -.0485 .I000 .0286 -.0619 .0238 -.0567 .0213 -.0534 .1500 .0321 -.0697 .0267 -.0639 .0237 -.0603 .2000 .0355 -.0753 .0296 -.0687 .O264 -.0650 .2500 .0391 -.0796 .0324 -.0726 .0288 -.0685 .3000 .0424 -.0827 .0353 -.0756 .0313 -.0711 .3500 .0448 -.0846 .0375 -.0774 .0333 -.0727 .4000 .0464 -.0854 .0389 -.0779 .0344 -.0732 .5000 .0447 -.0822 .0379 -.0751 .0349 -.0716 .6000 .0363 -.0694 .0334 -.0668 .0323 -.0652 .7000 .0268 -.0537 .0246 -.0539 .0235 -.0538 .9000 .0089 -.0177 .0082 -.0178 .0078 -.0176 1.0000 .0005 -.0005 .0006 -.0006 .0010 -.0010
17.262 at 15.354 c, cm . . 23.843 20.838 9.538 at 33.381
LER/c . 0.0226 0.0188 0.0155
19 TABLE III.- VERTICAL-TAIL AIRFOIL ORDINATES
Vertical-tail water line at stations - 23.90 27.62 30.97 39.37 62.01 x/c Z Z =_z +-- :Lz +z ,de C C C C +_ 0 0 0 0 0 0 0 .0017 .0051 .0025 .0067 .0066 .0060 .0060 .0034 .0069 .0050 .0092 .0090 .0083 .0083 .0050 .0083 .0075 .0109 .0108 .0099 .0099 .0084 .0104 .0100 .0124 .0122 .0112 .0112 .0167 .0139 .0125 .0136 .0132 .0126 .0126 .0334 .0190 .0250 .0184 .0181 .0173 .0173 .0501 .0233 .0500 .0253 .0248 .0237 .0237 .0668 .0270 .0750 .0311 .0304 .0285 .0285 .1002 .0325 .1000 .O358 .0351 .0328 .0328 .1336 .0366 .1500 .0428 .0423 .0396 .0396 .1670 .0397 .2000 .O48O .0475 .0448 .0448 .2004 .0419 .2500 .0515 .0513 .0486 .0486 .2338 .0433 .3000 .0539 .0540 .0514 .0514 .2672 .0440 .3500 .0551 .0556 .0530 .0530 .2859 .0441 .4000 .0554 .0563 .0539 .6144 .0441 .4250 .0554 .0563 I .0539 .O539 .6326 .0437 .4500 .0554 .0563 .0539 .0539 .6660 .0415 .50OO .0552 .0555 .0531 .0531 .6994 .0383 .5500 .0532 .0534 .0516 .0516 .7328 .0346 .6000 .0496 .0499 .0490 .0490 .7662 .0306 .6500 .0449 .0452 .0450 .0450 .7996 .0263 .7000 .0393 .0394 .0394 .0394 .8330 .0220 .7500 .0329 .0328 .0328 .0328 1.0000 .0003 1.0000 .0004 .0004 .0005 .0013
c, cm . . . 48.330 33.670 29.174 23.139 9.764
LER/c . . 0.0060 0.0100 0.0100 0.0083 0.0083
2O TABLE IV.- DATA FOR CALCULATIONOF PROFILE DRAGCOEFFICIENTS
(a) Wetted areas and reference lengths
Model component Wettedarea, Reference length, am2 am
Fuselage andwing-root flap track fairings . . . 10 257.9 171.20 Wing ...... 5 733.9 19.50 Horizontal tails ...... 1 820.7 14.30 Vertical tail ...... 1 651.2 21.66 Pylons ...... 213.2 34.45 Nacelles ...... 1 479.2 34.54 Outboard flap-track fairing ...... 261.7 17.60
(b) Reynolds number per meter, wind-tunnel stagnation temperature, and calculated profile drag coefficients
Values of CD, p for configurations - M R Tt, OK BW 1 BWlHV BWlHVN 1 BWlHVNIT
0.550 10.62 x 106 301 0.01508 0.01929 0.02093 0.02113 .625 11.35 307 .01481 .01894 .02055 .02074 .725 11.94 315 .01453 .01858 .02016 .02035 .750 12.13 323 .775 12.25 322 .01439 .01839 .01996 .02015 .800 12.40 316 .825 12.50 328 .850 12.76 325 .01416 .01810 .01964 .01983
21 c
r_ c_
v_
E
c_
E
w_
_o E
c
LL.
22 L-68-988
Figure 2.- Photographs of configuration BW1HVN1T with sting strut mounted from the bottom. L-68-987
23 \ \
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[ t Lacelle kicatioH q 'l [; ' Inboartt flap-lrack IJ_ring Asl)ec_ ralJo _ 4[ ChJtb,_ar_flap-_rack lairiny i Spar, 175 5'_ ! Incidence angle L D_he_ral a n_jle i
(c) Planform geometry of basic win 9 l
Figure 3.- Continued.
26 ro
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29 I Area 1114.27 cm2 17,71 ii Aspect ralio 4. DO _. Taper ratio .40
33.38
14. 30
16.12
2384
9.54 8.54
J i' I I I
(g) Horizontal tail.
Figure 3.- Continued.
3O -j
_ c_)o
_ l_ v _ L_
c_ m
31 L-OS-IIZ9
Fi9ure 4.- Photographs of configuration BW2HVN2 with sting strut mo,Jnted from the bottom. L-68-1130
32 L-68-1174
Figure 5.- Photographs of configuation BW3HVN 2 with sting strut mounted from the bottom. L-68-1172
33 Moment relerence center. _ J /
"'-. 4.LL__,__.=_ ==--; WL 12, 50 F-- i- r Sting strut 35.3e
i
/' / Center of model rotatior, f (a) Sting strut through bottom of m_'de!
Tunnel center line \
"\,\
",,,\ '\
WL 12.50
Cb) Sting strut lhroJgh topof model plus dlJF II?
Model station I30.1 Tunnel sta. 4084
Center of Frlodel rotation --,
/ f / 35.36 L SUng strut
Tunnel sta 414_
Ic) Sting strut through top of model
Figure 6.- Sketch of support systems. All dimensions are in centimeters.
34 (a) Sting strut through top with dummy on the bottom. L-68-781
(b) Sting strut through the top. L-68- 782
Figure 7.- Photographs of configuation BWI showing alternate mounting systems.
35 .002
O---£ ----C, <3_-O-£_--C_----{ ---(1 _ M M O.55 0.55 0
O---O---C <>O-O-£ -OH3--< ---q_--< .625 .625 0 ,.-
C---O---C O-O----C-C -C_S.-_-C-.__( >__.< .725 .725 0
O---O---C <>fL ___r£ _%-C---_ ___< >___q .75 CA_i .75 0
.775 .775 0
C"_-0 ""--C 'O-C'__ _-_-C '-'C ---<. _--{ ) .so .80 0
) 6""-'--'0-----_3 <> i 0 """_ ---C CJ-C -0-0 _'S _ _" < --.--6 .85 .85 0 I --4 --2 0 2 4 6 8 ew, deg Figure 8.- Variation of nacelle internal axial forces with wing angle of attack. Internal axial force referred along fuselage center line. 36 M=0.55, 0.625, 0.725 -7 __Z__ ACD,ov .002 __ 0 / .002 -- Mo75j_ °: ACD,ov 0 J i -.002 .002 M = 0._____ I ACD, ov 0 -.002 - 0.85./_ _ / .04 .825, 0.85 0.80 ACL, ov .02 M =0.55, 0.625, -_ --"_ 0.725, 0.75, 0.775 0 .02 M =0.55, 0.625, 0.725, 0.75, 0.775, 0._ AOm, ov 0 - -.02 , I , J , I , , I , J , J , J , I , I , I , I --4 --2 0 2 4 6 8 aw, deg Figure 9.- Variation of incremental strut-interference forces with wing angle of attack. Extrapolation of measured data is represented by dashed portion of the curves. 3'7 0 BW I [] BWIHV 24 20 16 2 aw, deg 12 0 8 -2 L 4 -4 D 0 .20 -4 .16 -8 .12 -12 .08 .04 C m .04 .03 0 - .04 C D .02 .01 - .08 -.12 0.4 -.2 0 .2 .4 .6 .8 -.4 -.2 0 .2 .4 .6 .8 CL C L (a) M = 0.55. Figure tO.- Longitudinal aerodynamic characteristics for configurations BW1 and BWlHV with 6h = -0.50 . Transition fixed. 38 0 BW I [] BWlHV 6 24 20! 4 ,61 2 ih "w, deg i2 ! 0 21l:1:, !l _ _ f 8 :xtxt -2 L D - T- TT T_ii_" ! ii 4 X -4 i i I I,_ ! i i! i ! 0 : ! : .20 : ! -4 i11:_ ;q: I. : .16 22!!_ ,]i !i .,.+..,+.x .12 -8 !;11 -12 _ : Cm .04 .04 ii !i:i::ii!iii_iiiii!ii-¢-I !!!:i:;_ ii.i:!: iii _i:!i i!i ¸ i[!i !!!i ii!! ![! :1:1i :]: ...... ]!!i":__iixi', , ....!iii FT_TT_ 0 .03 :_:¢::-!!i] 2!1"' !i!! !iiiiiiiiiiliiiiiiiiiiiS" :iiii¢iiiiii,,, ,,,,, , ii!i lit =;_'_ . ::x:] , : CD .02 ...._' !]!i _i]i !]! !]! ii i ...... i::ii! i i::ix:.i .Oi i¢i .... !ii_iiil2::i:2::..! .... ¢i;iii:':ii; i;ili ""_'i ;!! ...... [Jill !iii!i!!!!i:iii_!!!_!i_-T:_;i!iT::!iii;ii ili!!i!i!iiiii i: _o.£ 0 .2 .4 .6 .8 0 .2 CL CL {b) M = 0.625. Figure 10.- Continued. 39 0 BW I [] BWlHV 2O :!<1i:_::!: i:¸_i • i i::i¸ _ii 16 I I : !1 i:ikl 2 2 12 !!i ! ':i: ...1 .... :: !..... i k 8 iiiti<: L ,:_i:< f !r ; i -Ti 'T ' I /'1 [ _.r:1 1 ....l.i 1 :i L TT-- : ? i? i::i i i 0 .06 _:_ iii!lil ill...... -:i+.: :. E::: • =_Tt::= .05 i! ii :iil:ii :{ 5: iiiii!il iiiii[iii .04 ...... _- 7_ ¸ GD .03 =I= L .02 r !! ...... ol i i : o --.4 .4 C L 0 L (c) M : 0.725. Figure I0.- Continued. 4O 0 BWl [] BWlHV 8 iil:_ _r ...... 1 6 ii:!i!i r::f:::T 2 [T!_FT "'-+ _'" • C,w, deg 0 ;u:b::_ I::: ..... ii! --_- _ .... -...... 1 1¸¸¸ i:;gi:!::F::liiil...... !.... ...... ! :i i : I --4, _ .2O !!Siiiiiii---? ...... _"" ::TI :r ...... i ..... ::: -:it: --t_- ,I, i.iL 06 .... :i!t .... i ::.i:: :! l:.:Jid: • : ii_i:!i :"if"" _"f"" !FT!! • i i' 1 T-'iT'" ['..iLl. '_LI ::::I::::qi!:Jill .05 _:iq_:;i;...... : '! [iiii:? :' : i_sli I_! .... _'-_-'-f---l-iiii!i ...... , Xi it{- _ .... _:iill. i:!it!:_ilillii • .... i : : :! :!: , _d4...... i;;;ill iii!iiiiiliiii[ _:h., ltr:1 I .04 :/_ii:;::i;i;i:_E ;:i!_i_:i!!!!!!!1!!!! C m • 04 ::i) I t t.... I :i_ IL,_i ;i ...... : i: fix:_...... * i!ii;ii:!li_!il_:'::""!!:!:!!iil;lli!;_iiii!!_!_ i,b*,-f--;ii:; !75i!_!i==::I:: : :j: I t" ' [ [ f!:": ! CD .03 s_...... ;:-i.i,.!:!!!!:t...... ii!il;!:il!i_l"...... iii 0 ::)_:iiii:f!!ti!l:::. i:t , _: i ...._.... iiii:,ilT:!...... i_ffiiM! ,, ,,_ : . I h.t t i i : ! !: !2 [ t -.04 ;::::1::: I t | I ] :: ,; ,: .... ] :: , , b_-t- I: _: b-_ ..... ;T_-: b i!!i ..... !!!!!_-_f!:fb:_ __P'i': kK!_..[:: : _:ii_ii ...... _i i i !::::!:::.1:::I: _ i I i!iiFi_ ! - .08 I-r! !-!! J _ I ] .01 "" :-!:: .iLiiii:; _iii_}-:i_:iii:::Liliiii : : : : : i!! _ !: !-' i_i._i:i_iit!_iI!iilill !!:il!! I!i! I 'k :ii:F: "k :l k F!ii:i" t ii:ili_i -.12 ..... % '-.i.... 0 .2 .4 .6 .8 _.4 -.2 0 .2 .4 .6 .8 CL CL (d) M = 0.75. Figure 10.- Continued. 41 0 BWI [] BWIHV 6 "!:i ....7-i iii.t:-_ :i:: _1 _4q: _i: _ 2 i:i!:i•i 4 i-T : I ::T" : i i I ...... ] :i , _ I i J i : i1: ! _w, deg L L : x 4:: D .... 0 i¸!1114 £ [z£ -2 xt !LI¸_ e i i i !_i $..L_::_[± - - 4''1' _ !::i .2Ll , ::i :: _4 -!71 ';}: ; [ I -4 : t::_IX -] -;: " it : t-; :TIT: ,K:_: .07 .iiiii!i ZXlilZ T:Ti .06 '' :L.I::I: i!!it!:!! iF ..+ ...... _T i:i .... -:-IT-:- .: i : : I .05 _,i_ : [: ! _=.i:i :i! -4 - : i:ii: :: ::L i_::i! C m Z :[: CD " I i I ! .03 xxbx --++ -i .... .02O !i 31;;_i qTT_ ::xl ::: :q :1 : : : iilLI _lx: : i t-l-r , 'i ...... :r: : _ _21 .; :!::: _1 x 0 H:!q [ --.4 --° 0 ,a CL CL (e) M = 0.775. Figure 10.- Continued. 42 0 BWl [] BWIHV i_iitiiii _::: I"-7- .... _ !iii:_:_ :_'!t i ¸ ! ::i:]i:!_ 4 ...... __ 2 L aw' deg D T ...... 0 __ _ __ 0 L : ii " I -2 i i i : t i ...... i:Ig: TT__-_- -4 .08 :., ...... ,_:Lf!i!i_ h_i_- .... i'll _ ! i T _ iLi- :1 • -_-+-- iiii _ --_-- .o7 :_::!::::iiiii!ii::_:!:_::i, I :I _ 'i t i i i i : ' i i ::i :i : t:t.:1._ .o_::l,,iiiii:,:i:-: iiii_fff I i .05 :::::::::::::::::::::#:i_i::: ::: i::i!i::!!!_ii: i _t./: CD .04 ;l;i!;i;;ili.::L_ i i[i ]: G m :iii: 'fJ Pt:I -P:i t:i:l:i: =+" )i}i)ii_.::4.: iili ":: iili_ Fil:! ! .03 :i::!:i::_ i!': _i !i!i_!!! ...... _) ...... F:F:k :x[: i: ! !_ i .o, Ii ii!ill ::::::::::::::::::::: ;:::i;:: :::$::: '77:::: *-_-- ....i:.i i: ;:1 i 1_ -.4 -.2 0 .4 .6 .8 -. -.2 0 .2 .4 .6 .8 CL CL (f) M = 0.80. Figure 10.- Continued. 43 o BWI iq BWIHV 12 L /, i_i_iii_ii_i_iiiliiiiiiiiii:!iiiiii: D aw, deg iZ2_ZZ -4 -2 -8 -4 !!!i!!!!iii!- __,__ _ax_: -4- fiiti;i :::q':: ?i!i_ :21_1!L1!!121 ....t'" [TTT: ....+...... !: CD -4... _ .... t- :'[TY " !- :T ::!El: I d::T 0 .2 CL CL (g) M = 0.825. Figure 10.- Continued. 44 0 BWl [] BWIHV :::ii_:i_ :i_ i_::i::::!::i_ ! i¸:! !_!!_Li!/ ...... : !_ ii I IFI .:.:I..:_tiiiii L -' : ...... !ii -- 0 :--:!:,: i:i ! :!.:::::f!! _:_::_I_ _:_!_:_i._: D r:'::r:-: --+-- F.[_!!': ...... ii i:'# r::]: ] ; : : -! !1:!!_'!!!!!:_:- ih: ...... !i_ ::-:_i:_ .... _:'_ ...... ift: _;ifi_ ffi;:ffii i!ii i i _ i _ ! _,_,: ...... i.... _ i , b i ! i G" ' =: .... :L.. :i:2 .,..,..,C_CI ...... :!'? ':_i _t_ t;iiT ::{'iii ...... i IF-II iFIiil F[!]!III ii_i I t:iii ,:_l !;iR_X!iti .08 !!_i'! t: _ ]! .o7,t!!!?_' !iiil_ -.+_._ _+__ , + + ._.,i ?1 i ? h!.:! _ .o5 :ii!i;ii,iiilIil i ii!_ t CD ...... i:ii il-_; :T::I !!i!i_: : :!:t_u!l O m !i_ii!i !!i:ii! F!i !{£!:i: !!:! iii:ii: : -t i ::UT =::!: ili;i;: i] .I.!.] : : I! =-..!!i!i! :!"f :t'"fi -i!H ! ! ;: =1 ::_ 4--4- 2" '! o '_ !IHiiii ; ; :i: ; ; ;ql;;;I -.4 -.2 0 .2 .4 .6 .8 -.2 0 CL CL (h) M = 0.85. Figure 10.- Concluded. 45 Transition 0 Fixed r-, Free 6 20 i_i!L i 16 ::7!! q _ :! : 4 _ii! iiiti!i! 7i17i t2 !<- 2 _ [:i ...... ! ':i iii_ i ' i i ikl :! ! aw, deg UT.= [i+ 0 8 ; I ! ! t _ :: i; iii :i!igi:IT!_:ti:::i:::i : ! :i L !7!!! D ,.q- t : 4 ._L,, -2 < i: i _i::-j.!i it i iii liit _!iiit:::iFil:.i | i_ TT_U! iii ill ii: ...... 0 dL)ii -4 -i 'I iii li: _i_i!!A -4 .20 U_ 7111' _ i!iii:i:ilili -8 •Io !!_ i_i i l : .08 _ ..... [ G m .04 .... i 7 CD o I -il -.O4 _ -.os ]: :! CL CL (a) M = 0.55. Figure ll.- Longitudinal aerodynamic characteristics for configuration BWiHVN1T ,_,ith 5h = -0.50 . Transition fixed and tree. 46 20 ilil i! !ili i ...... ii it_:i-.-,_.... ::-_ , , ! _ ] ! ii ; + 16 _------_ iiti!!: i _!!_i : i : ----: [ ! ' F--7_ _1,4_ 12 i i =::ix:: ;:::;;::: !:i_.i:::11121:1 7 , ...... 1: r::;::x; ...... 8 ;: g_ii::. ....iiil!i;i:.,_i -4 i i- _m iiL: i i_ _ : ! i : !il ;: =t .... ', ...... i ! ---,---4 ; , Cm .04 .....: i )ix CD i ! :iL:i{' 1 i --+...._ ...... _-.h--' list - .... !7 ...... _-_F+_ _!_t .4 .6 .2 .4 .6 .8 CL CL (b) M : 0.625. Figure 11.- Continued. 47 Tronsifion 0 Fixed Free 2O i_i-i!!!!!!! !:::: iii i_ii_iI :J:;! :!. !! !!!!!!!i_ 16 4 :i q i i ii:!!_iii :_1:"1 1: - 12 2 iiii i!ii;_:i- _:" 'iii!iiii_ aw, deg iiii_iiiiiii!iill2:![[211 i_i_i 0 • m::: 1:;: _I:! il i_!iii!:ii!!i! L i ill D _1__ 4 -2 iiLiiii 141::: iil] L: }i Ll .....!iil...... ! I I il;iiii: iii!i'iii...... _ .:.-it:.... _-7:! -!1_!!!!Ti:" 0 -4 rT -4 iiilfilli!iiii!i !!_!I!!!!!!!!I!1!! -8 i= =..:!.i ! i;iii; .I 2 _l:_1:.. LL:ILL: ÷_+__ !!!!1!!!!!!_i!_!! .o8 _ _- : :r C,m _N ,_! !! !.. " "ii!!i fi "-T: :!i!! 0 D 0 ; i;i;;; i [: -.o4 _:: i ! ! I: i=i:: -.o8 i i ][: LI2! ,:_ f,i= io_ 1:, 1:1i _)rT- 0 - -i,_ -.2 o .2 .4 .6 .8 OL (c} M = 0.725, Figure If.- Continued. 48 Transition 0 Fixed Free 16 : :i: : t: 12 4 ::I: ! l . I ! i i ...... !-!- , -:!. _!:: : i :: deg ii :f l • L I !"l "_: ; OlW, D ! !::i G', , -q-'- :7 • 0 :!:i -- 2 :::q=: I- -4 ! t : ::: - 4 ...... : ::_::: ....::: ::::: :. ; :.! -8 ::' i iii .20'_ii_i_: ::[:h i ! 1 • I 6 iii[iii . L_ :i:lx: ]: I I :, iiii iii ...... !iiii: !: ...... ::x .12 iiiii_i _f ...... :!!:i :-::::: ::'T_ i lill .o8 :;:!:1; [ , :!: Crn .04 i_iIiill CD .q .... • r- _: .i I i i ::i_:'i :T_t r:l'i : k: !!!!!!i!!N...... : ...... ! ] .... !_:jr_ !:!!i!! .i!!i'i :,::: i! : -.o4 ii!iii; i!!i!i! - ::::.:; :m::: t :H t - : : :i!i ii ii:Iiil-,...... : -•o8 _iiii_i Z'Z_T: }i ..... , .2 .4 ,6 .8 _°d .4 .6 .8 CL CL (d) M = 0.75. Figure 11.- Continued. 49 Tronsition 0 Fixed Free 16 12 6 4 L aw, deg D 2 0 -4 -8 .07 .06 .05 CD .04 -.08 -.12 .8 -.4 - 2 0 .2 .4 .6 .8 OL. CL (e) M -- 0.775. Figure 11.- Continued. 5O Tronsition Fixed Im Free 8 12 :i_: i!"¸ i i ' i -b:_ -T i;F!FII ii!i_iL ii: _i_.l_i_T¸_ ] i!ii!ii!!i!iiii ii : ¸1:1¸ _t:!tlt:;: 4 !i;!i_ii_iiiii!! 4 L D aw, deg 0 2 iiiiiiiiiii!liii!!_i_ii,i -m: ,, -4 ii_ii!i_iLii;ijil;_l:ii, 0 -8 -2 -4 .07 .06 .O5 CD .04 0 m 0 .2 .4 .6 .8 CL CL (f) M = 0.80. Figure ]1.- Continued. 51 Tronsition 0 Fixed Free 6 8 4 L 2 0 D aw, deg 0 -4 -2 ...... ::i_iiii:,i i:.:_:.:...... _q<_:T_i -8 4 .09 , :-_,_'_:,:_, .O8 F_ [-- i il];i_ ' • t .... < 1i! ' I ...... :,il ::!!._!i::__iii;...... Y:': ...... : _i:: ...... ;__ 1 ..... :!: .08 ...... :!i! !!! CD 0 m .04 .....:i! !!;:t .... : t ] !,!!!!_;: ::=:!ii, .03 :L iit t=1 -.04 t _i_: t .02 ] !< i' j :=l'. i -.os_:ilt i :i; ill .0[ ::i!I t_:i: ' !)_ i i: i: i :i , i .4 .6 .8 -.4 -.2 0 .2 .4 .6 .8 CL CL (g) M : 0.825. Figure ii.- Continued. 52 Transition Fixed Free 8_iiiiiiii!}ii!Ii! 4 iiii i!ii]iiii L 0 D -4 -8 i:: ::].... .o9ii!!ii_ i:::ii,_ .o8 !H_:_ =F _:!i!i!!i:i i'!i_iNq ":_ml IF:F;': CD .05 !_i_i!_l!i!ii2!i!! .04 _:_*':_:_...... ;;;i'_i'_;i: titi:ii !!i!tii!ii_!4i+ glili_!ii:!}ii:: L£::: _!_!i!NI :[;it:if'. .OI , !!212!!21::j:::: i H :I :: : :::I : -.4 -.2 .2 .4 .6 .8 CL CL Ih) M = 0.85. Figure 1]..- Concluded. 53 Transition A Fixed G Free 2O : Li L:tI:!_ 16 4 ::::::::::::::::::::::::_i_i:i_i!ii iiiiiiiii iii iii_iiii'iiil --L- ::i 12 2 aw, deg 8 0 L D 4 -2 O -4 -4 -8 Cm iii!ii!FiiilF!iiii::ii_i 0D _! t:i:' -.o_ ...... _ (/ !ii: F_'H! :!!: ill: I ? :!: iit:i ii: i I ! :il ;- : i: -.12 1: -A -.2 O .2 .4 .6 .8 0 L CL (a) M = 0.55. Figure 12.- Longitudinal aerodynamic characteristics for configuration BW1HVN1F _ith 8h = 0°" Transition fixed and free. 54 Transition a Fixed a Free 20 6 i i i i._ ! i 16 4 12 2 ew, deg 8 0 :i_iiiiiil.].L L I D 4 -2 0 -4 -4 -8 G m CD .01 0 .2 -.2 0 .2 .4 .6 .8 GL CL (b) M = 0.625. Figure 12.- Continued. 55 Transition A Fixed _q Free 2O 8 ili!i: TTT _ Ii iiilK 6 16 : : : t:::: 4 12 i: i: :!.... _ii!:i ? ! !t!i :iiiiiiil itiiii!iiiliiiiiiiiiiii!l! :i!i i:i,•. :! : i ! _w, deg 2 8 i]il L I?: :i !: i;_ill . ; D s_ff ::_ff:!÷ 4 0 iiiitiiil iiilliii "]T T! i i .... 0 -2 , ._ !!;!_t!]!_:!::.f! - _i! • _i_iiil!. !ili_i _:'_: iT- : ; :!T :::::::.: :.t: -4 -4 ...... I I. t; 1 -8 .16 ii-_;i; ;÷!--__!;i;; :: i!; ...... _. I::iIi':i:::_..,.i-_i:#i_!:! .12 :!i: iii! -q !! ! !!:i ?:!i:: :Z .08 i: :i!it :!) %_:i_ 52!_i2i !!i!:: ¢ .O4 .04 .:iiii_ '!i_ 0 m ! t!i i!ii:! :t:!i i!ff_ i-i ii_ -iiiiii ,:iti_-ii;4;;- !;ii{_:,! .O3 !fi:11 ii!it:_ ::::x:: CD ?iil :?2:;:_-.;::.:;.: ..... ,.. : 22;22 i: i!ili )_iii!E ::::i:::!!!!I_,:_:ll}; !::!!_...... h!_ :,iii_: .O2 ::i!::: -,04 1:: .... !--- ::- : ii{ :ilE$1! :i .01 :i !i: - .08 :;'," .... i'- -:-'/= ' [ i ii :i::!i !!i:! !!::I:!:111111::!!!l!!!_!!!Ii';!i:_i':....__i_i i :..ii:i!i I -.12 0.4 .8 --.4 .2 .4 .6 .8 OL. OL {¢} M = 0.72_. Figure 12.- Continued. 56 Tronsition /x Fixed Free 16 8 i_il ' -i4-- i ...... iiiF []_ -- ! ii_ i- 0 ! .... 0 _i iii i_iii!/i : I::: Li:_ ¸¸;¸ ili_!i_ _i_ii_i -4 -2 :x: 1 ;i:t ;il x i::_ :i!!!;!! ::ii -8 -4 :!: ]! .16 !12 :x::: _:_ L!IL _.x:x ::!ilx ii::l:i:i =:4 ...... _...: ;: [ I ,,14,:: .12 i_iIiii :i;/ :<:!71 :::: : 7['_ :[;: ::r:ttr :: -< _].£_ _TT iili :::x: .08 ;i!!i_i :x :::: ::: .L..] i]] is_]::j i !i i2i!il .... I ]] .04 !!;Iiill C D C a _iii!ili 21:ili:.: k 0 ;_';]:_ _::iiiii -.04 iifi!il T.t:Z[ :lit .Oi -.08 _!_i_ !q::F N ilii=:= 1:i.._ _ .2 .4 .6 .8 -.2 0 .2 .4 .6 .8 CL CL (d) M = 0.75. Figure t2.- Continued. 5'7 Tronsifion A Fixed n Free i ,2 1 :Ii + :.:_.i:;i..::,:::4_i_-ii' I .....2 12...... --t. . i:! i l :!:!2: -:: i_ 8 !! i 6 L 4 4 D [ i ;: '_ ._i i:: "!::;'" ] , -]:+T -4 0 --8 -2 _..__...... _----+...... :...... i:-T-:_7::I:: .O8 - 4 .O7 .16 .06 . I 2 .05 i .08 0D .04 .O4 0 m .03 .O2 -.04 ..... .OI 0 --,4 -.2 0 .2 .4 .6 .8 • -. -;_ o .2 .4 .e .8 OL OL (el M = 0.77.5. Figure 12.- Continued. 58 Tronsition A Fixed D Free 12 10 8 6 L D 0 4 aw, deg 2 0 .07 CD i t ¸_ 0 .2 -,2 0 .2 .4 .6 .8 CL CL (f) M = 0.80. Figure 12.- Continued. 59 Transition A Fixed Free 8 6 i_!;ii_i_;,_• _...... L 0 4 D !_iiiii! -4 <, aw, deg 2 -8 0 .O9 -2 ...... !i !: ...... ! i,_.;_ii !L: :_, :i .O8 -4 .O7 .O6 .O5 CD .O4 C m .O3 .O2 .01 0 -.2 0 .2 .4 .6 .8 OL CL (g) M = 0.825. Figure 12.- Continued. 6O Transition /x Fixed D Free IO L O D -4 4 aw, deg -8 0 -2 -4 CD C m 0 .2 .4 .6 .8 CL CL (h) M = 0.85. Figure 12.- ConducJed. 61 Tronsifion l& Fixed o Free 20 16 12 2 ew, deg L D 0 4' -4 .12 i i ! , ili i-!: -8 .o4 t'_! _: Cm 0 .05 -.(?4 _1 ! CD -._6 : _o.4t-.2 Iii o .2:_ .4il ii.6 .s -.4 -.2 0 .2 .4 .6 .8 CL CL (a) M = 0.55. Figure t3.- Longitudinalaerodynamiccharacteristicslor configurationBW1HVNlfwifh 6h = 0.5°. Transitionfixed and free. 62 Tronsifion [x Fixed o Free 20 6 16 4 12 2 _'w' deg 8 0 L D 4 -2 0 -4 -4 -8 .05 .! i ! i i _ : i i i _ ' i i i ; : i i '[ _._t i, C m CD It -.2 0 .2 .4 .6 .8 CL CL (b) M = 0.627. Figure ]3.- Continued. 63 Tronsilion Fixed Free i k 12 'i .¸ i -! i-T: 8 t _/,r'i i I ! aw, deg L D 4 T " '_' i I t !I i | 0 - 2 ! _ :_'i •i -I ! i i -I ! ...._---I:-F-_- , / ! ! ,: i :i i i i i -4 i 4Lii ], _il i L_ i it L -8 .12 .O6 .O8 .O5 .O4 .O4 Cm 0 -.04 CD .03 .O2 - .08 -.12 -,1_6.4 --'.2 0 .2 .4 .6 .8 CL CL (c) M = 0.725. Figure 13.- Continued. 64 I2 lili i: i:iii L ii 3!_!II-!!!II!i!...... T 4 iiiiii_ii::ii i_ii_ I L___::: _;;i Zi 4 ;:: • i2 _ , _.____,L...'.._21 L ' I ' " " L i: i i i : i , , i .04 : C m ! ' i ;7I-T-- 0 .:'_._i...:a I; -.04 !i:':_iiili i _ !F: -.o_ ::L ii : t-! ....:T-i_,i i __ _t i t I -.12 iq. : I _ T--: .... : : I t ii x:Ttr_:: :{ : I 17i .8 -.4 -.2 0 .2 .4 .6 .8 CL CL (d) M = O25. Figure 13.- Continued. 65 Tronsifion Fixed © Free 16 I0[---- 12 8 4--i-- i :' 8 3 L D 4 aw, deg 0 ...... :...... t iy -4 0 I i i -8 -- 2 " .O8 .O7 , :,....!_i_! i:__ I i .O6 .os i:_< t:i:: I , ' E .05 .O4- 0 £D .04 Gm .O3 - .04 :,_-_,_o:: _ _ hi !:i :i .... .O2 - .08 ...... 4--4...... i i i , i - i i :. _ ' _ ; ! i i I i: I.LI ...._:--.::i_:.i -.12 .01 :I i _, i , :._L :: : ::TF_,-: ...... T_,-T: T- -.16 -% -.2 o .2 .4 .6 ._ --.4 -.2 0 .2 .4 .6 .8 CL CL (e) M = 0.775. Figure 13.- Continued. 66 Tronsition I& Fixed o Free 12 !_ Jo I-:7- ! 4 L D 0 4 -4 _ili_¸_i_!_ii_i_ _ _!_ -8 _ 0 -2 -4 CD C,m CL CL 6'/ Transition h. Fixed o Free 8 8 4 6 L -- 0 : :! D -4 : i i ii i I :! ii. ! _ aw, deg 2 -8 #i:__ ili_l_?t_ii_!_ .O9 .O8 -4 .07 .12 .O6 .O8 .O5 .O4 CD .O4 Cm 0 .O3 -.04 .O2 - .08 i!_ ii_!!:_¸'¸_¸¸_,__i ¸¸_ _i_:_ _:;_:,_ : i.LLi:: :_LL- .01 -.12 0 -.16 --.4 -.2 .2 .4 .6 .8 -.4 -.2 0 .2 A .6 .8 CL CL (g) M = 0.825. Figure 13.- Continued. 68 Tronsition r,, Fixed 0 Free 8 I0 !.... il, _ii[!i i:i .i.i_:i!_ 4 [, i 8 i_ I i L i ': I i i D 0 I ! : I ' i'i !;!; i'i....! _w,ae_ 'i;: -_ i ['_ ..... _: _ _d i _ _ ' i i -8 ii_ili_iiiiI?!_i!i , ...ii_ 2 _._ .10 :i_!i:i :ii.:) "i.... _ ;:L!IiL -2 -.L- .09 ii_i _!_' .08 -4 ._+ .... i _i.. I. _ :['!T iil'_ .07 i!: :::: :: .12 _TI:I ,O8 0D .o_ ] .... ,04 IiTFTT' .04 Cm 0 : :::::: .03 12[ :::: d [ I ] - .04 11;[;i ili _ii!;11i_;i I ...u- .02 - .08 -.I 2 _!il_i:!!ii!¸¸ i I I I t .01 _.+.. 'El t T!_I:" :j:.: 0 ' i r --,2 0 ,2 ,8 -,4 -,2 0 .2 .4 .6 .8 CL CL (h) M = 0.85. Figure 13.- Concluded. 69 <> BW I HVN I d_ BW2HVN 2 o BW3HVN 2 2O 16 12 4 L D -4 -8 6 ] 2 -12 Cm .04 0 CD OL CL (a) _, = 0.55. Figure 14.- Longitudinalaerodynamiccharacteristicsfor configurationsBWlHVN1, BW2HVN,2, and BW3HVN2with 5h = -0.50. Transitionfixed. '70 0 BW I HVN I r_ BW2HVN 2 (D BW3HVN 2 20 6 i _ t i i ! --T---- 16 4 J2 2 i¸ ew, deg 0 L D -2 1 -4 -4 .20 -8 -12 .05 .04 G m .03 CD .02 .01 04 -.2 0 .2 .4 .6 .8 CL CL (b) M = 0.625. Figure i4.- Continued. 71 <>BWIHVNI BW2HVN2 E] BW3HVN2 ;6 6_ 12 4 8 2 _w, deg 4 0 L D 0 -2 i-i--!- -4 -4 -8 .2O -t2 .16 .06 .I .O8 .o_ 2 _I-il _....!-_-i::___:_ _b_-t Cm ,04 .o.Lk:_::i:l,li I [:_¢t; CD 0 .o_! i'_I _l!l_i__,_Ii - .04 .o, ii]!": ii_i!=.===-_i[ - .08 -.12 _o.4i:_-.2:"_[-f:_lo ...... 2 [i .4"-i_ .6::ii .....8 -,4 -,2 0 .2 ,4 .6 .8 GL CL (c) M : 0.72.5. Figure 14.- Continued. 72 0 BW I HVN I BW2HVN 2 E] BW3HVN 2 I _,l 8 !T!-ii---iiiii-i lift L D ! ! !;;LL:: I.:_l _' deg 02 ) ii;T- _ T -!i iF: I -4 .08 ...... 20 iiiii_. •i t: . i !!_!!!:-i!!i_-1 hil !_i: ! , _1 !!ii !iii! : .07 ,i _ !!1 ....i---_ z .,6 i!¸ iiiil i...... i ! ..... _1 i_ i i i I ! i i .o8 :!'_ _-i .05 TTtlii_i:' I :_i;;_t!:?_?:_i:::i:::i...... _:ill_;i i -_ iliit ,i. ii!it f_T "_: !..! i;ii Om .04 " i 4..... i _L _ i 0 ! i : : . , , • : : i : : [ '::T7_ q .... ¢'1 V .02 _ :L:_::_' i:;:;iiiii - .04 : " i : i;4:4 ; : ::.:i, . : ] - .08 :i:: I 'ffl: I -.40] _t_.:2z_f!;_li Oi -L2.4 ±.2' o .2 .4 .6 .8 CL CL (d) M = 0.75. Figure 14.- Continued. "/3 BWIHVN I BW2HVN 2 BWsHVN 2 16 Io[:-Ti- _,...... = t2 6 L 4 4 D aw, deg i:i ¸ !i _.....i,i,_ i i_i o [ / -4 -8 .09_ -- 4 L ...... I .O8 :i i _ i ,20 !_i _ _: _ i i!i i ..... i...... :-- i i i i i i! .O7 i! .,6 i, il i_ i Ii. [i. .O6 .O5 CD .O4 C,m .O5 i ...... l J_i ' ---i...... +-.'-}-iq .O2 -.04 ' i i i _:_i _ _ ! t[:,_.:::::::i::]:t::£2:1:,,_i=2:].,2:1\_,, _...... : L : .01 - .08 0 -,12 ...... -.4 -.2 0 .2 .4 .6 .8 -,4 - ,2 0 .2 .4 .6 .8 CL CL (e) M : 0.775. Figure 14.- Continue& '74 12 it ¸ , 4 L i,. D 7 i 0 -4 -8 .09 .08 ....+ • _ .07 i .06 ,12 .05 ,08 CD .04 C, m ,04 .05 0 .02 - .04 ,01 - .08 0 -.12 --.4 -.2 0 .2 .4 .6 .8 -.2 0 .2 .4 .6 .8 CL CL If) M = 0.80. Figure 14.- Continued. 75 0 BW I HVN I 13 BW2HVN 2 0 BWsHVN 2 12 8 _:t i Lt:I_:!. !:!::::_::g: 8 __:L_LI I t_:: : :÷!"F-- _ _ ; I t. _ :i: £ iIi + 4 L D _w, deg 2 -4 -8 -- 4 _ ! i CD O m .2 CL CL (g) M = 0.825. Figure 14.- Continued. ?6 o BWlHVN I BW2HVN 2 (3 BWsHVN 2 8 "i!!iiii:il!i_:,Z 22122 ._ ...... i • _ .i!: ....i¸¸¸:_L. ; _2 ] [ :. : ! i 4 i !iF _ ! _ . :_: : ::_: , ÷-_-*'_ '-:-i -L i!i:_ _ [ 1 i i _w, deg 2 :[ i...... i!ii,-¸ •! ,....i!i,i, -8 o :i : I I i _:i •_o _- :L, !"!i!:! '_ -+ 7[Z .... ii!iii .09 t -4 : i _ i '! !!i_:!l':!I!:_, :1::,i::i:_ikk_ .20, .08 : _ ::i !ii_:_i 1 : i I I ---_ r-: - .07 i: _ j_:. , [ I I! ! .16, ;,i ...... +.._ ::i:4 t .... • • L .06 i_ , .12 ! L i n;:; , i-i ;11;. i! i .08 :!!_¸ c_ .04 .03 ;!ii ::L!l:ii2!:: 0 "! ! ...... --t --_'! .02 ;;!;;;;::!:i::!i!! -.o4 :: ...... i : !11i T1...... ' • +---_- ÷.y, .! ;[ I t: : ! -.O8 !: ,i _}ii i I 0 .2 .4 .6 .8 CL C L {h) M : 0.85. Figure ].4.- Concluded. 7,? BW I BWIHV 8WIHVN I BW I HVNIT CL a CmCL .12 .O8 .O4 Cm, o 0 , , . . (a) Lilt and stability characteristics. Figure 15.- Effect on the aerodynamic characteristics of adding horizontal and vertical tails, nacelles and pylons, and flap-track fairings with 6h = -0.5°. Transition fixed. 78 BW I ..... BW I HV BWlHVN I .... BWI HVNIT -:- 1--:"-v CD CDat CL=0.2 i ' .08 ¸¸:¸¸[,! ! : , //// / .0 7 .06 / J_, .05 L CD .04 .05 iJ 11" .02 .01 C D aT CL = 0.3 _._ _ - . CD at CL=0.5 _ii ! , 0 .5 .6 ,7 .8 .9 .5 .6 .7 .8 .9 M (b) Drag characteristics. Figure 15.- Continued. '/9 BW I BW I HV 8W I HVN I BW I HVN iT ,,!._!i_t:iirf hi,....!i...... :iii! r TI i:l,!i ....!, Hlil i!I_i_t_i!': _=;! i 4h4 CL for ii!ti, _1!!:',| ,4 (L/D)max 11i l!!l :T/-!!!i 24 _T-_:: Z _-- r::, : _ :: .I ! ,if'': ! :!_ ili!i': 2o @_!ii ::! t_:_i_ !iiii!;l:i: :!: ?:i_:i!:i::iliii:'!iil ii_iii i (LiD)max 'ii hl ],:_?! !iJ!i!i_;i ./:i..LI :i/Jii_ :_: ol...... :I? _t?:-?:._!_.,_t,,,? ,5 ,6 .7 ,8 .9 M (c) Lift-drag-ratio characteristics. Figure 15.- Concluded. 8O BW I 8WIHV .024 8W IHVN I BW IHVN I T .020 .Oi 6 CO, p .01 2 .008 .004 0 0 AC O at CD, min [] Z_CD at CL=O.30 .008 .tiHl!itJtil!lill.f:it:.t_ilIilhl_l :_i_:ili:i. .OO4 ACD, N I 7;tit!i!':!i!il:_!i'_f'dNl!ilIlillit::itil,itittilittit,itit'_ttI_tlif::iP;f_i if:4:;i!i!i!it!i!im ': ': _ _ _: _itiftitNtftit t_t:tttthhHitttit,;t_httit_,_tititNittifftiiiti i!i _ _ ,_[ 0 ...... ::_ii:;i_,[_!_!/i! _r,,_r777iriiiT!717i-77l'i7 ,i' i:it;i!ii t l!_!17 !! ![7tJ/i: .9 M Figure 16.- Computedprofile dragcoefficientsandmeasuredincrementaldragcoefficientsfor the nacellesand tails. 81 _n, deg - 0.5 0 .5 .16 .12 CL ,08 01 04 CmCL Cm_o • 6 .7 .8 .9 M (at Lift and stability characteristics. Figure 17.- Effect on the aerodynamic characteristics of horizontal-tail deflection _or configuration BW1HVN1T. Transition fixed. 82 GD 0 - r ...... ...... CD (b) Drag characteristics. Figure 17.- Continued. 83 o ¢' IsO "o do CO o I ca ¢...) i - g _k_kl'_'_7"it_...... _I"r'_--!_''_.. CO o 0 " 84 .040 .056 .052 .028 .024 CD, trim .020 .016 .012 .OO8 .004 _!i_i :,iI_ iiii 0 .I ,2 .3 .4 .5 .6 CL, trim (d) Trimmed drag polars. Symbols represent trim points. Figure ]7.- Concluded. 85 .16' ,12 CL .08 O/ _04 0 0 CmCL --.2 --.3 --.4 0 --°1 Cm, o --,2 --,5 .5 .6 .7 .8 ,9 M {a) Lift and stability characteristics. Figure 18.- Effecton the aerodynamiccharacteristicsol fixing transition lor configurationBWlHVNtT. 86 Transition Fixed Free _m_ o _!_i`_ii_!_;i_!_!_ii_i_i_i_i_!_!_!!_1_Uii_ii_!_i!I_ii_iii _:%_:_i_i:_I_i_i_:_i_/_i_i!_i_!i_._i_i_i_i_#_J_._iii_!ii!i: .,2 ii _ _=°° i!i iftif!f!itillttitittt!iJtiltlllt!llil!l!ilillt ltttillililltlil#ftittftlilltit!tfittIl_i_111!!!i!!!!!!!!!i!!! ,v.*,_,_H_flli_!!_i[i!_!i_lHHl_ii!_i_i_i_i_!Hi_J_!i_i_i_I_!ij_i_!_i_#_i_i_i_!i!_!_!_H_'' _IIIIiI1111111111_!itli!!iI;H !I1: ililtt!il ' t _t_t_t:_:_t_ _i_tiiit ittt,fit:f,ttf:t::fltii t!fll! !! [!:!!!!1tlt:t!1ti Cm_o iiiiiii_i_I t_l_tt:l_littiltttftittftit!fttt:ttitittt!ittt!itttif!tf_f_!_t_i_! i!i_ _F ili! !itittiI!_it_miiiiiiiiiiiiiiil;iii',i;i_ftifif!ttil',tti!i!tf!:,ltf!ttl!ii_ ._i ilH_!f! tH:IHV. itltt_FIT!F_ttl .04 :iii iiii _i;iiii;iIH[ t I!itlt!ft_tiltttiftHHttif!lt:'t#ffiltItit':ftit!f!tii_ili_lll t,itti,i.tt;i.tt_f_t,iY,ifit_f!it..ftfItttf_f_ttft,tfifittft_tfftttlt_J[Ll,it!ltit::__ _ii_i!_¢_!I_!i_iii_i_iIi_!i_!_!_!_i_`_!i_`_`_ii!i_i_!_!i_i_!i _!!i!!i iit;'ii!i Ii ! lil_ llsltHitil!il:i I':HIIIII_,_!i_i_I!i_i_[i_i_li_Hl_!*_!_!l_f!_i_Ii_!_I_i_¢IIi_!H_i_il!_i_i_!_i!iIUi!Y ! i!il o _i;;i;_iiiiii! l tti l ii tt I_ itlilll_lil!iiI'l!i!}!!!!!!!! "= _!_II -o.5_;ii_i_!_i_iIiH_!_H_!_i!_i_!_i_!_iI!_i_!_I!_H_iI_!_!i_i_!_!_i!i°!ii_!_!_i_i_I_Ii_!I!Iii_!_i_i_i_i_iI_i_!_i_I!_!_i_i_!i!_[_ ii i_ittititlfiitti! h- _il!llillI!IIillI!llil!llilillHill[_Hllil!ll_l_II_i_IIllllilllliiIHI!llililliliHi_ilili!i!tf!i!l!#tti _iiiilili_illi .08ill!liI!l!iI!liiI!iiiiHilll_lilllli_liilliL_ll_IIIiI_iIll_il_I_iIl_II_!l!il!llilil!llililll!II!_if!lt,lit!II_l,t _l!llilil i_ll!l_lllJ,lll_ Cm_ o _!i!_i!i_1i]i_i_!i_i_iIi[_i_I_I_i_I_i_i_!i_!I!__HI_!_i_i_i_i _0 4 i!iili]iilHill_.._ :' • I_I!IIihlllINillll_N_ll _i,/_' :[HIHHIHIIIH ! Ii! {ii!li!i!ll!i!llilillilillllil!I!ilitlIl_liIII!illt/lll!iI!IIl_i_l_IIii_,IIiIi':-- I_!!!!!ll! ii!li!lll_.lli_illlli!i illil!illlillllH llili I i I _ !llilillllillIIIillii!l__ I II_.41__ _ _iiiiiiii_iiil .5 .6 .7 .B .9 M (b) Pitching moment at zero lift Cm,o. Figure 18.- Continued. 8'7 Transition Fixed .05 ...... T-_ ...... : ...... Free ' ! f.f i[ • f r _ , , '+ ...... _--+- :' "1- ...... : ..... ,L__ [ : ] ,' i i j _ , .03 ....+-._--:-?4 ...-:._...... i _ .+ _ I/...... i . + , i i : I : i I ll/ I I 0 2 [ --_ I-.p:_= [ = _ .: = [ T : I: - , _I _ 4 I .ol:-:::':-Zi-.-L-: i : .i. ' ...._:--f---i-:.....I CD,m_o . .09 ...... ! .08 ..... .07 ...... .06 ...... CD .04 ...... .6 .7 .8 .9 M Ic) Dragcharacteristicswith 6h = -0.5 °. Figure18.- Continued. 88 i .!: _: iii ¸ ::4::: :: : :: .... i ..... !...... i ..... ..... "-'---d-'- 1 ..... ' i. :...... i ...... i:: _: i i:--.13 : i i ! i t ;'; '!= '" i r r _ . . 14-) II o _D o4 _0 _t o 0 E E _u_u_ 3 o I ._J I I i I .o_ !_ _ii ..... t lI_ _ _i _ _...... I I ...../7 i/ i . _:: _..... I : _"...... I- --- : I ': r/. ' I .... i A I I i_t ! _ _ ...... '.:i i- i--_i:!: _- - .., ...... _ _.__..-_.. _!...- _-- I : i ! i i! :! I . [ i _. I i !i I:_ co _. _ _ 0 o o" 0,1 x X 0 _ E E _..I _- 0 --J ._I 8g BWtHVN I BW2HVN 2 BW3HVN 2 .16 .12 CL .08 a .04 0 0 --.I --,2 CmCL --.5 --.4 i .12t Gin+ 0 I 0 .5 .6 .7 .8 9 M (a) Lift and stability characteristics. Figure 19.- Effect on the aerodynamic characteristics of wing modifications with 6h = -0.5o. Transition fixed. 9O .o5_ .... _ ...... 1 _...... ' ._. !:i !:_:!:, L i..... / _ ' _ i _ : ; _ _ _ .04!:-, : t i. _ i--i. _. L-1-i, : _ : ! CD ,08 .0 7 .06 / .05 CO .04: / / / - I , //V / 1 .0:5 .O2 i .... .01 GD of CL = 0.3 L.,_.L. 0.-5 ' B6 .7 .8 .9 M M Ib) Drag characteristics. Figure 19.- Continued. 91 I111 __-+-+-- i i i i. i i , i_1 i i ! 1!11! _040,1 zzz III an cn 0B :I:MII I ,..2 E l = g m N_ = I = g I o l ;111i I i g! ;llr4 ttt- tO NHi fill 2 I I I I t_ 0 o Lib 0 03 92 -+_ _+_ O " -- CXl C',,.l ZZZ L -I--I--I- E -I.- C) COl I_ m .5 0 .._I o '--+_ ....i CO O " x x 0 t_ B E ._I NASA-Langley, 19'70 -- 2 L-7141 93