1、IIIII IIIIII 1JI!llllrmlllrlillllll11 III IIIIII 3 117600501 9352 NASA Contractor Report 165843 NASA-CR-165843 19850021606 APPLICATION OF AN AERODYNAMIC ANALYSIS METHOD INCLUDING ATTAINABLE THRUST ESTIMATES TO LOW SPEED LEADING-EDGE FLAP DESIGN FOR SUPERSONIC CRUISE VEHICLES Harry W. Carlson tlDRARV
2、 COpy “Jp 1 3 1982 KENTRON INTERNATIONAL, INC. Hampton Technlca1 Center LANGLEY RESEARCH CENTER LlBRPRY. NASA HAMPTON, VIRGINIA an LTV company Hampton, Virginia 23666 Contract NASl-16000 t1arch 1982 NI5f National Aeronautics and Space Administration Langley Research Center Hampton. Virginia 23665 “
3、fOR EARLY DOMESTIC DISSEMINATION / “kecause of Its significant early commercial potentI, this mfOrmatlon, which has been developed under a US Gov ernment program, IS bemg disseminated within the United States In advance of general publication This information may be duplicated and used by the recIpi
4、ent With the ex press limitation that It not be published Release of this information to other domestic parties by the recIpient shall be made subject to these limitations. Foreign release may be made only With prior NASA ap proval and appropriate export licenses This legeld shall be marked on any r
5、eproduction of this mformatlon m whole or m part -, ReView for general release tlarch 31, 1985 , 1111111111111 1111 111111111111111 1111111111111 NF01332 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-SUMMARY A study of low speed leading-edge flap d
6、esign for supersonic cruise vehicles has been conducted. Wings with flaps were analyzed with the aid of a newly developed subsonic wing program which provides estimates of attainable leadlng-edqe thrust. Results indicate that the thrust actually attainable can have a slgnificant influence on the des
7、ign and that the resultant flaps can be smaller and simpler than those resulting from more conventional approaches. INTRODUCTION , The highly-swept low-aspect-ratio wings which permit high levels of aerodynamic efficiency at supersonic cruise conditions present serious problems in the low speed flig
8、ht regime. One of these problems is the achievement of a sufficlent1y high lift coefficient to perm1t safe terminal area speed at an angle of attack which does not limit pilot visibility. The required lift coefflcients can be generated at acceptable angles of attack through use of trailing-edge flap
9、s. Unfortunately, for conventional supersonic cruise designs with wing-mounted engines and outboard ailerons, only a small portion of the trailing-edge span may be used for this purpose. Thus, large flap deflections are required to generate the additional 11ft, and drag penalties may be excessive. P
10、roperly designed 1ead1ng-edge flaps can bring about significant improvements in the aerodynamic eff1ciency without reduction of the lift coefficient or increase in the associated angle of attack. As reported in reference 1, significant progress has been made in improvement of the aerodynamic efficie
11、ncy of 1eading- and trailing-edge flaps for superson1C cruise conf1gurations. The convent1onal approach to leading-edge flap desiqn has been to place segmented flaps on all of the wing area ahead of the front wing spar and to conduct wind-tunnel tests to determ1ne optimum deflections. A somewhat dif
12、ferent approach to the leading-edge flap design problem is the subject of this paper. The concept is based on the observation that the pr1mary purpose of the flap system is the achievement of an aerodynamic efficiency comparable to that which could be atta1ned with full theoretical leading edge thru
13、st. Accordingly, the new approach first attempts to assertain the local degree of achievement of leading edge thrust for the basic wing. Then, as required in a design by iteration process, local geometry changes 1n the form of leading edge flaps to compensate for the loss of thrust are introduced. T
14、hus, for portions of the wing 1ead1ng-edge where full theoretical thrust may be anticipated no flaps need be employed, and for the remainder of the leading-edge the flap chord and deflection angles may be limited to values just sufficient to restore the efficiency losses due to the failure to develo
15、p full leading-edge thrust. The use of the computer program of reference 2 in the estimation of attainable leading-edge thrust and in the prediction of the aerodynamic characteristics of flap configurations is shown in comparisons with experimental data for a generic Sllperson1C transport model. Fur
16、ther application of the computer program in an iterative design mode is illustrated in a sample problem - the definition of flap geometry for a typical supersonic transport in landing approach. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-AR b cA
17、cN cr cD CL CA CN Co CL CL a LI L2 TI T2 M Y s S R 2 SYMBOLS wing aspect ratlo, b2/S wing span section axial force coefflcient section normal force coefficient section thrust coefficient section drag coefficient section lift coefficient axial force coefficient normal force coefficient drag coefficie
18、nt 1 i ft coefti ci ent lift curve slope, dCL/da designation of leading-edge flaps designation of tral1ing-edge flaps Mach number 1 atera1 distance from Wl nq centerline sucti on parameter, wing reference area Reynol ds number angle of attack CL tan CL tan (CL/CL ) - CD a (CL/CL ) 2 - CL /(nAR) a le
19、ading-edge flap def1ectlon angle, positlve for 1eading-edqe down trailing-edge flap deflection angle, positive for trai11ng edqe down local angle of wing surface at the leading-edge relative to the free stream direction, includes basic wing camber and leading-edge flap deflection Provided by IHSNot
20、for ResaleNo reproduction or networking permitted without license from IHS-,-,-Subscripts n measured in a plane perpend1cular to the h1nge line s measured in a plane parallel to the free stream DISCUSSION Assessment of Computer Program Applicability The computer program of reference 2 which provides
21、 estimates of attainable thrust for wings at subsonic speeds is based on a planar solution of linearized theory equations. To study the applicability of the program to the present problem, comparisons of program results with previously unpublished data from tests conducted in the Langley Research Ce
22、nter V/STOL Tunnel have been made, and are shown in figure 1. The wind-tunnel model employed in these tests is particularly appropriate for this purpose. It represents a M = 2.7 cruise vehicle, but for simplicity only the wing and fuselage are represented in the model and the wing has no twist and c
23、amber. The low speed test conditions are M .28 and R 5.7 X 106 The program has inherent limitat10ns in the accuracy of flap planform modeling due to the wing element grid system employed. Although the spanwise pos1tion of the flap edges could only be approximated, the flap areas were matched by comp
24、ensating changes in the flap chord. In figure l(a), the program results are compared w1th data for the basic flat wing. There is good agreement between the theory and experiment for the full range of angles of attack “and lift coeff1cients. The axial force correlation is particularly slgnificant sin
25、ce it shows an appreciable degree of achievement of lead1ng-edge thrust. The normal force curve shows evidence of the presence of vortex lift, Wh1Ch 1S also accurately est1mated by the program. Figures l(b) to l(d) show similar correlations for a series of leading-edge flap deflections with the trai
26、ling-edge flap deflect10n fixed at 10. Both trailing-edge flaps see sketch in figure l(a) were set at 10. The correlations are not as good as for the undeflected case, but there is still a reasonably good prediction of the 11ft-drag polar. Figures l(e) to l(g) show correlations for a series of trail
27、ing-edge deflec tions with lead1ng-edge flap deflections maintained at 30. For this 30 leading edge flap deflection, axial and normal force predictions are poor. There are, however, compensating effects so that the lift-drag relationships are given reason ably well in the CL = .4 to CL = .8 range. T
28、he program is seen to underesti-mate the amount of lead1ng edge thrust and overestimate the normal force. The ability of the program to assess trends may be examined with the aid of figure 2. Here data from figure 1 1S shown as a function of leading edge and trailing edge deflection angles. The suct
29、ion parameter s is defined as in refer ence 1 to be a measure of drag relative to the limits for fully attached and fully separated flow. These results indicate that, despite some inaccuracies in the absolute values predicted, the program may be used in a design process. 3 Provided by IHSNot for Res
30、aleNo reproduction or networking permitted without license from IHS-,-,-The Des1gn Problem The configuration of Table I has been taken as an example for application of various flap designs (see reference 3 for an explanation of the format used for the geometric description glven in Table I). This 1S
31、 a wing-fuselage-vertical tail conf1guration with a twisted and cambered wing des1gned for CL = 0.10 at f1 = 2.7. Landing approach des1gn cond1tions have been chosen as: M = .25 R = 160 X 106 CL = .55 a = 8 Two trailing-edge flaps on either side of the airplane (between the fuselage and the inboard
32、engine, and between the inboard and the outboard eng1ne) are fixed 1n planform but may be deflected as necessary (the same angle for both). It is assumed that trailing-edge devices for the rema1nder of the wing will be employed as ailerons for roll control and will be unavailable for use in generati
33、ng lift. Conventional Design Approach As a base-l1ne reference, convent10nal lead1ng-edge flaps simllar to those treated in reference 1 have been analyzed. In that reference the test results 1nd1cated that a uniform deflectlon along the entlre leadlng-edge performed as well as, if not better than, a
34、ny other deflect10n schedule included in the tests. Accord1ngly, the convent1onal flap analysis w11l be simpl1fled by the assumptlon of a constant deflection over the whole of the leading-edge. Results of the analysis are summarized in figure 3. The slmplif1cation of one deflection angle for the tra
35、illng-edge flaps and one deflection angle for the leadlng-edge flaps permits the program results to be presented in the form of a contour map. Suction parameters at CL = .55 and angles of attack corresponding to CL = .55 are shown by the contour lines as a function of the lead1ng- and tra11ing-edge
36、deflect10n angles. According to the map, the optimum performance of the flap configuration subject to the limitat10n of a eriment TheOfl - No thrust - Full thrust .24 .16 Co .08 Attainable thrust o .4 .8 15 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS
37、-,-,-.04 -.04 .8 .4 o-.4 _1O-O.L-ILO_-J20 a,deg Figure 1. - Continued 16 8 -8 -8 T ,n - T,I,n - T ,2,n -8 -8 L,n - L,I,n - L,2,n o Experiment Tf“tC!)fY - Clo thrust - Full thrust Attoinoble thrust Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.08 I
38、 / o-+- -.04 .8 .4 CN 0 -.4 -10 o a,deg , ,f“ 8 -8 -8 T,n - Ttl,n - T ,2,n 8 -8 -8 L,n - L,I,n - L,2,n o Experiment Theory - No thrust - Full thrust Attainable thrust .24 .16 Co .08 (g) 8T n= 40 t 8L = 30 , .n Figure 1. - Concluded 17 Provided by IHSNot for ResaleNo reproduction or networking permit
39、ted without license from IHS-,-,- 00 a,deg for Cl =.55 Suction parameter, s for Cl =.55 8 -8 -8 T ,n - T,I,n - T ,2,n 8 -8 -8 l,n - l,l,n - l,2,n 8l,n =300 8 = 10 IS T,n 12 8 4 0 1.0 .8 .sL 0 0 0 L 0 0 0 .4 .2 00 10 20 30 40 0 8T,n ,deg 8l,n ,deg Figure 2. - Comparison of program suction parameters
40、with experimental data. 6 M = .28. R = 5.75 x 10 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,- lO 40r-1 I I I 8L,s ,deg 30 20 10 00 10 20 8T,s ,deg 30 40 8 =8 =8 T,s T,I,s T.2,s 8 -8 -8 L,s - L,I,s - L,2,s Suction parameter, s Figure 3. - Suction
41、parameters and angles of attack at CL = .55 for full span leading edge flap system. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-20 - No thrust - FuJI thrust Attainable thrust -.04 .2 O- y b/2 Figure 4. - Spanwise distribution of forces on the bas
42、ic camber surface at the design lift coefficient. CL = .55. a = 10.48. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,- No thrust - Full thrust Attainable thrust .12 .08 ct cas EL /: .04 _ - v-.“, o:=d .08 .04 -_ . -Or-=-=-=- -.04 -.08 -.12 .8 .6 .4
43、.2 -;:- r-O:=-1 O- .2 .4 .6 y b/2 Figure 5. - Spanwise distribution of forces on the basic camber surface at the design angle of attack. a = 8. CL = .425. 21 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-N N 40 30 8L,s ,deg 20 00 10 20 8T,S ,deg I
44、I I I I 30 8 =8 =8 T,s T,l,s T,2,s 8 -8 -8 L,s - L,l,s - L,2,s .6 Suction parameter, s I / I I 40 Figure 6. - Suction parameters and angles of attack at CL = .55 for partial span leading edge flap system. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,
45、-,-8 =8 =8 =240 T,s T,I,s T,2,s 8L =8L I =8L 2 =350 ,s , ,s “s - No thrust - Full thrust CA Attainable thrust O-+-.04 .8 .4 CN OTT-+-.4 -10 o 10 20 a,deg .24 .16 o -.4 o .4 Figure 7. - Program aerodynamic forces for partial span leading edge flaps. .8 23 Provided by IHSNot for ResaleNo reproduction
46、or networking permitted without license from IHS-,-,-24 .08k .04 Ct cos EL 0 16 .12 .08 cA .04 8 =8 =8 =240 T,s T,l,s T,2,s 8L =8L I =8L 2 = 350 ,s “s “s - No thrust - Full thrust Attainable thrust . O-r- -.04 -.08 .8 .6 .4 .2 .4 .6 .8 1.0 Y b/2 Figure 8. - Spanwise distribution of forces for partia
47、l span leading edge flaps at the design condition. CL = .55. a = 8. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8- = 8 I = 8 = 20 ! s T. ,5 T,2,s .4 8 -8 -8 Lps - L,lts - L,2,s L b/2 o ;:;:;:“,:= 0 to .5 .567 .633 -1.2 -1.6 I-I-I-I-I o 10 20 30 40 8L,s ,deg .700 .767 .833 .900 .967 Figure 9. - Variation of section drag-due-to-lift factor with deflection of partial span l
