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MEMORANDU M
LOW - SPEED INVESTIGATION OF THE EFFECTS OF HO RI ZON TAL-TAIL
AREA AND WING SWEEP ON THE STATIC LONGITUDINAL STABILITY
AND CONTROL CHARACTERISTICS OF AN AIRPLANE
"
,
CONFIGURATION HAVING TAI L SURFACES
OUTBOARD 0 F THE WING T IP S
By William C . Hayes, Jr., a nd William C . Sl eem an, Jr.
OTS PRICE
Langley Research Cen t er
Langley Field , Va .
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NATIONAL AERONAUTICS A ND
SPACE ADMINISTRATION
WASHINGTON
June 1959
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w 9 CONF IDENTIAL B NATIONAL AERONAUTICS AND SPACE ADMINISTRATION MEMORANDUM 6-II-59L LOW-SPEED INVESTIGATION OF THE EFFECTS OF HORIZONTAL-TAlL AREA AND WING SWEEP ON _HE STATIC LONGITUDINAL STABILITY AND CONTROL CHARACTERISTICS OF AN AIRPLANE CONFIGURATION HAVING TAIL SURFACES OUTBOARD OF THE WING TIPS* By William C. Hayes, Jr., and William C. Sleeman, Jr.
SUMMARY A low-speed investigation was conducted in the Langley 300-MPH _L 7- by 10-foot tunnel to determine the static longitudinal stability and control characteristics of a model having tail surfaces located outboard u and rearward of the wing tips. The wing of the model had an aspect ratio of 1.00 and could be adjusted to give leading-edge sweep angles of 60 ° , 65 °, and 70 ° . Four horizontal tails of different size, in which the area varied from i0 to 27.8 percent of the wing area, were used in the tests. Very brief tests were conducted to assess the directional stability characteristics of the model for angles of attack up to 30 o.
The test results indicated that the outboard horizontal tail was an effective pitch control over the test angle-of-attack range; however, for the basic model with the 70 ° swept wing, there was an appreciable loss in longitudinal stability for lift coefficients above approxi- mately 0.70. A reduction in wing sweep angle to 60 ° improved the lon- gitudinal stability at high lift coefficients and reduced the variation of stability with lift coefficient throughout the test angle-of-attack range.
Directional stability of the model was high throughout the test angle-of-attack range, with the stability at an angle of attack of 30 ° greater than at low angles of attack.
Title, Unclassified.
CONFIDENTIAL w _ @ w @ • • • •I_ CONFIDENTIAL INTRODUCTION Airplane design trends which result primarily from the quest for high lift-drag ratios at supersonic speeds have given rise to trouble- some stability problems at moderate and high angles of attack. A pos- sible airplane configuration which may alleviate the stability problems encountered on many high-speed configurations yet may maintain attrac- tive performance at supersonic speeds has been suggested. The basic concepts and some supporting experimental results are presented in ref- erence i for an airplane arrangement having the horizontal- and vertical- tail surfaces mounted on slender bodies and located outboard and rear- ward of the wing tips.
Tests of an outboard-tail configuration at Mach numbers from 2.50 to 3.51 reported in reference 2 indicated that relatively high values of lift-drag ratio for trimmed conditions could be obtained on an arrange- ment which also had good directional stability characteristics through the angle-of-attack range. The results of reference 2 and subsonic data of reference i, however, indicated possible problems of longitudinal stability at moderate and high angles of attack. The present low-speed investigation was therefore undertaken to study the longitudinal stabil- ity characteristics of a simplified model, geometrically similar in most respects to the model of reference 2, and to explore means for improving its longitudinal stability at high angles of attack.
Longitudinal stability characteristics of the basic and modified model were obtained over an angle-of-attack range up to approximately 30 °.
Effects on stability of both wing sweep and horizontal-tail area were studied for wing leading-edge sweep angles of 60 ° , 65 ° , and 70 ° and for horizontal-tail areas that were I0, 15, 20, and 27.8 percent of the wing area.
SYMBOLS The system of axes used in this investigation is shown in figure I together with an indication of the positive direction of forces, moments, and angular displacements. The moment reference center was located 52.45 inches from the nose and corresponded to the 50-percent mean aero- dynamic chord of the 70 ° swept wing. All coefficients presented herein are based on the area, span, and mean aerodynamic chord of the composite plan form of the 70 ° swept wing and largest horizontal tail. The symbols used are defined as follows: CONFIDENTIAL ....... : :-- : :-- :'.
v_ qvv w _ _ I we we • • _o _o Lift lift coefficient,
C L
qS Drag drag coefficient,
CD
qS Lateral force lateral-force coefficient,
Cy
qS rolling-moment coefficient, Rolling moment
C_
qSb pitching-moment coefficient, Pitching moment C m qS_ yawing-moment coefficient, Yawing moment Cn qSb reference span of wing plus span of largest horizontal tail, 4.00 ft reference mean aerodynamic chord based on combination of wing and largest horizontal tail, 2.16 ft mean aerodynamic chord of horizontal tail _h mean aerodynamic chord of wing alone _w free-stream dynamic pressure, ib/sq ft q S reference area of basic wing plus area of largest horizontal tail, 6.96 sq ft plan-form area of horizontal tail
%
plan-formarea of wing alone s_ CL angle of attack of fuselage center line, deg angle of sideslip, deg A angle of sweepback of wing leading edge, deg deflection of horizontal tail with respect to fuselage center 5 h line, deg C ONF IDENTIA L
L
MODEL ANDAPPARATUS
The model (figs. 2 to 5) used in this investigation consisted of
mahoganycentral and outboard bodies and aluminum wing and tail surfaces.
The basic wing was swept 70° at the leading edge and had an aspect ratio
of 1.00, a taper ratio of 0.39, and a flat-plate airfoil section 0.017_
thick with a faired leading edge and beveled trailing edge. The basic
horizontal tail was swept 60° at the leading edge and had an area
27.8 percent of the basic wing area Sw. The wing leading-edge sweep
could be varied, whereas the wing span and area remained constant.
Geometric characteristics of each wing plan form used are presented in
table I. In addition to the variation of leadlng-edge sweep angle a
leading-edge extension of 4 inches in the free-stream direction could be
attached to the basic wing. All horizontal tails had the sameplan form,
but area ratios Sh/Sw = 0.278, 0.20, 0.15, and 0.i0. The samehori-
zontal tails were used with all wings; however, for a few tests, the
horizontal tails were movedforward 4 inches when the configuration with
the 4-inch leading-edge extension was tested. All horizontal tails could
be deflected about a hinge axis through the 25-percent meanaerodynamic
chord of the horizontal tail and perpendicular to the center line of the
outboard body. Vertical tails which were identical to the basic hori-
zontal tail were used in all tests.
TESTS AND CORRECTIONS The present investigation was conducted in the Langley 300-MPH 7- by 10-foot tunnel at a dynamic pressure of 57.5 pounds per square foot which corresponded to an airspeed of about 150 miles per hour. The test conditions produced a test Reynolds number of approximately 3.0 × 106 based on the mean aerodynamic chord of the combined basic wing plus basic horizontal tail. The angle-of-attack range was from -4 ° to approxi- mately 30 ° . Two tests were made through the angle-of-attack range with mm the angle of sideslip at T ° and -5 °.
CONFIDENTIAL .......... .: --- --- -Z'. i'" i :'" i': ,up ! v v I • oB .v • v CONFIDENTIAL Blockage corrections as computed by the method of reference 3 were applied to the free-stream dynamic pressure. Jet-boundary corrections as computed by the method of reference 4, with the combined basic wing and horizontal tail considered as a single lifting surface, were added to the angles of attack and drag coefficients. Corrections for tunnel buoyancy effects were also applied. No corrections have been applied to account for the base drag of the model fuselage.
DISCUSSION Inasmuch as the tail surfaces were assumed to contribute a positive increment of lift to the total airplane lift in trimmed supersonic flight, the aerodynamic coefficients for all configurations are based on the com- posite wing area of the basic configuration; that is, the area of the 70 ° swept wing plus the area of the largest horizontal tail (Sh/S w = 0.278)._ This configuration is similar in plan form to the supersonic model of ref- erence 2. The moment reference location was at the 50-percent mean aero- dynamic chord of the wing of the basic configuration (A = 70o), that is, 52.45 inches from the fuselage nose, and was assumed to be a reasonable airplane center-of-gravity location. Since the position of the outboard bodies remained fixed with respect to the fuselage for all configurations, it was believed that this moment reference location would be satisfactory for all tests.
Effect of Horizontal-Tail Deflection The effects of horizontal-tail deflection on the aerodynamic char- acteristics of the model are presented in figure 6. These results show a progressive decrease in lift coefficient as the horizontal-tail deflec- tion is varied from 0 ° to -12 ° at a given angle of attack; however, the increases in drag coefficient associated with the larger negative deflec- tion at a given lift coefficient are not so pronounced. This character- istic is consistent with the results of references 1 and 2 and occurs as a result of the field of upflow in the vicinity of the horizontal tail.
The pitching-moment characteristics presented in figure 6 show that the horizontal tail provides an appreciable stability contribution at low and moderate angles of attack. In addition the tail is an effective longitudinal control to at least 30 ° angle of attack as is indicated by the fact that the stabilizer effectiveness _Cm/_5 h generally is about the same at the highest angles of attack as at 0 o. The pitching-moment curve for the stabilizer setting of 0 ° indicates a loss in stability at the highest angles of attack, which is not noted at the other stabilizer angles. Apparently this loss in stability results from tail stall since CONFIDENTIAL - w w .
...... _ w ii ,D • • ill oi_ • _ • • • • ii iii i i i ii i iii i i ii i i i 6 CONFIDENTIAL the expected relief of the condition at negative tail deflections is verified by the remaining pitching-moment curves.
A significant reduction in longitudinal stability occurs for all stabilizer settings at angles of attack above about 16 ° (C L = 0.70) and is believed to be a result of changes in the flow angularity at the tail as high angles of attack are approached. A loss in effective upwash at the tails would be expected as the wing-tlp vortex moves up and inboard relative to the wing chord plane as the angle of attack is increased.
The problem of possible stability losses due to decreases in upwash at L high angles of attack for this configuration prompted the present inves- tigation of effects of horlzontal-tail area and wing sweep angle.
Effect of Horizontal-Tail Area The effects of variation of the horizontal-tall area for the 70 ° swept- wing model are shown in figure 7. Decreases in the horizontal-tail area were accompanied by reduced lift coefficients at given angles of attack and increased drag coefficients at a given lift coefficient; however, these coefficients were calculated from the reference area of the basic configuration; whereas the actual area decreased with decreasing tail area.
The expected reduction in longitudinal stability at low lift with decreasing tail area is shown in the results of figures 7 to 9. Of more importance, however, is the change in stability which occurred throughout the angle-of-attack range for each tail. A reduction in tail area had little effect on the variation of §tability with lift coefficient, and with the smallest tail (fig. 7) there was still a significant reduction in stability at angles of attack above about 15 ° . A reduction in tail area might be expected to eliminate the change in stability with lift coefficient by virtue of the fact that the tail-off pitching moments do not indicate significant stability loss at high angles of attack (fig. 6).
This improvement in stability variation with lift was not realized for the smaller tails, possibly because of the nature of the upwash variation across the tail span. The maximum local upwash might be expected near the tail root and the average upwash across the whole tail span, of course 3 would increase as the tail span (or area) is reduced. Thus, the change in upwash (resulting from displacement of the wing-tip vortex) as the angle of attack is increased is greater for the smaller tails and occurs at a lower angle of attack than for the larger tails. Therefore, the greater upwash changes apparently compensate for the smaller tail area so that improvements in the pitching-moment variation with lift were not realized with the smaller tails.
Z CONFIDENTIAL _. uw_ _-' oo • • q_w Ipo • g ,w O_B uo CONFIDENTIAL Effect of Wing Leading-Edge Sweep The results of figure 10(a) show fairly small effects of wing sweep for the configuration without the horizontal tail. Pitching-moment char- acteristics with the horizontal tail on show virtually no reduction in stability at low and moderate lift coefficients with decreasing wing sweep for a given tail size. At the higher angles of attack, decreasing wing sweep was accompanied by increased stability (fig. 10(b)). The pitching- moment results obtained with a stabilizer setting of -9 ° show that reduc- tion of wing sweep also effected slightly smaller stability changes with L angle of attack. (This stabilizer setting is chosen for discussion, inas- much as the stability loss shown for the 0 ° setting at high angles of attack was probably due to tail stall as mentioned previously.) Results with the 60 ° swept wing and horizontal tail (Sh/S w = 0.15, 8 h = -9 ° ) show very little change in stability with lift coefficient through the entire angle-of-attack range (fig. 9).
Effect of Extended Wing Leading Edge Characteristics of the model with the 70 ° swept leading edge with a constant 4-inch extension to the wing chord are presented in figure Ii for two tail areas _-ISh/Sw = 0.278 and_ 0.150) and in figure 12 for the large horizontal tail (Sh/S w = 0.278) moved 4 inches forward on the out- board body. No improvements in pitchlng-moment characteristics such as were obtained by using a lower wing sweep and smaller tall (fig. 9) were indicated. The results of figure 12 do show, however, that the charac- teristics of the basic model could be improved slightly by decreasing the wing aspect ratio and moving the horizontal tail forward.
Lateral Stability Derivatives Although the present study was c_icerned almost exclusively with problems of longitudinal stability and controlj a very brief evaluation of lateral stability was made for the basic configurations IA = 70 °, Sh/S w = 0.278, 8h = -9o) • These results are presented in figure 13 and show the model to be directionally stable throughout the angle-of- attack range (to 30 °) with the stability increasing markedly above an angle of attack of about I0 °. This trend has been noted in other outboard- tail models and is probably due primarily to the increased stability of the tail-off configuration (ref. i).
The variation of CZ_ with angle of attack indicates negative effec- tive dihedral at the lower angles of attack as a result of the contribution " CONFIDENTIAL i • of the horizontal tails at negative deflections and their long moment arms. A theoretical estimation of CZp at an angle of attack of 0 ° indicated positive increments of CZp contributed by the horizontal tail because of the tail sweep and aspect ratio when at a negative angle of attack (ref. 5) and because of the interference effect of the verti- cal tail (ref. 6). Inasmuch as the wings were at an angle of attack of 0 °, the only negative contribution to Czp was from the vertical tails.
The calculated value of CZp from the aforementioned considerations L was 0.00091 as compared with an experimental value of 0.00146 (fig. 13).
Although reference 6 pertains strictly to the conventional center tail assemblies, it was believed that the interference effect of the verti- cal tail on the horizontal tail would produce a small amount of positive rolling moment due to sideslip, which in turn, would be magnified by the long moment arm of the present model. At angles of attack above approxi- mately 12.5 o (the trim angle of attack for this configuration), C_p is negative.
CONCLUSIONS A low-speed investigation of effects of horizontal-tail size and wing sweep on the longitudinal characteristics of an airplane configura- tion having tail surfaces outboard of the wing tips indicated the fol- lowing results: i. The outboard horizontal tail was an effective pitch control over the test angle-of-attack range and would be expected to provide longitu- dinal trim for angles of attack up to 30 °.
2. Pitching-moment data for the 70 ° swept-wing model with a tall area of 0.278 of the wing area indicated an appreciable loss in longitu- dinal stability at a lift coefficient above approximately 0.70. Reduc- tions in horizontal-tail area from that of the basic model were not particularly effective in reducing the variation of longitudinal sta- bility with lift coefficient.
3- A progressive reduction In wing leading-edge sweep angle from 70 ° to 60 ° had little effect on the pitching-moment characteristics of the configuration without the horizontal tail; however, with the horizontal tail on, reductions in sweep angle increased the stability at high lift coefficients and decreased the variation of longitudinal stability with angle of attack. A configuration having very little change of stability with angle of attack was achieved by use of the 60 ° swept wing and small tails having an area 15 percent of the wing area and a tail deflection of -9 ° .
CONFIDENTIAL _. :........ • ..
_B CONFIDENTIAL 4. Directional stability of the model was high throughout the test angle-of-attack range, with the stability at an angle of attack of 30 ° greater than at low angles of attack.
Langley Research Center, National Aeronautics and Space Administration, Langley Field, Va., March 16, 1959.
REFERENCES I. Sleeman, William C., Jr.: Preliminary Study of Airplane Configura- tions Having Tail Surfaces Outboard of the Wing Tips. NACA RM L58B06, 1958.
2. Church, James D., Hayes, William C., Jr., and Sleeman, William C., Jr.: Investigation of the Aerodynamic Characteristics of an Airplane Configuration Having Tail Surfaces Outboard of the Wing Tips at Mach Numbers of 2.30, 2.97, and 3.51. NACA RM L58C25, 1958.
3. Herriot, John G.: Blockage Corrections for Three-Dimensional-Flow Closed-Throat Wind Tunnels, With Consideration of the Effect of Compressibility. NACA Hep. 995, 1950. (Supersedes NACA RM A7B28.)
4. Gillis, Clarence L., Polhamus, Edward C., and Gray, Joseph L., Jr.: Charts for Determining Jet-Boundary Corrections for Complete Models in 7- by lO-Foot Closed Rectangular Wind Tunnels. _R-L-123, 1945.
(Formerly NACA ARR LSG31.)
5. Polhamus, Edward C., and Sleeman, William C., Jr.: The Rolling Moment Due to Sideslip of Swept Wings at Subsonic and Transonic Speeds.
NACA RM L54LOI, 1955.
6. QueiJo, M. J., and Riley, Donald R.: Calculated Subsonic Span Loads and Resulting Stability Derivatives of Unswept and 45 ° Sweptback Tail Surfaces in Sideslip and in Steady Roll. NACA TN 3245, 1954.
" CONFIDENTIAL ir i0 CONFIDENTIAL L TABLE I GEOMETRIC CHARACTERISTICS OF W_NGS ALONE A, deg _w, in. Root chord_ in. Sw, sq in. Taper ratio Aspect ratio i. 000 6O 28.31 33.13 784.0o 0.691 1.000 65 28.76 36.Ol 784.o0 .555 1.000 40.23 784.oo .392 7O 29.79 .446 44.23 .875 70 33.53 895.45 All _ings had i/2-inch-thick flat-plate airfoil sections with faired leading edge and beveled trailing edge.
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W Lateral force P/fch/n_ momenf A I Ro///n_ mom_nf L/fl _ Yawing nTomenf
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Z Figure i.- System of axes. Arrows indicate positive direction of forces, moments, and angles.
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CONFIDENTIAL 13 i SIt A "60 ° oJ oO I Figure 3-- Sketches of wings tested. All dimensions in inches.
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C ,I Ji f- ,. L] i] tl f H -5 0 Figure 13.- The variation of the static lateral stability derivatives with angle of attack. A = 70°; Sh/S w = 0.278; _h = -9°" NASA 1 LangLey Field, Va. L-382 CONFIDENTIAL ... .
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