NASA TM x-246
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z TECHNICAL MEMORANDUM X -246
TRANSONIC AERODYNAMIC CHARACTERISTICS OF A MODEL
OF A PROPOSED SIX-ENGINE HULL-TYPE SEAPLANE
DESIGNED FOR SUPERSONIC FLIGHT
By Dewey E. Wornom
Langley Research Center Langley Field, Va.
Declassified March 15, 1961
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
WASHINGTON March 1960
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
TECHNICAL MEMORANDUM X-246
TRANSONIC AERODYNAMIC CHARACTERISTICS OF A MODEL
OF A PROPOSED SIX-ENGINE HULL-TYPE SEAPLANE
DESIGNED FOR SUPERSONIC FLIGHT
By Dewey E. Wornom
SUMMARY
Force tests of a model of a proposed six-engine hull-type seaplane were performed in the Langley 8-foot transonic pressure tunnel. The results of these t_sts have indicated that the model had a subsonic zero-lift drag coefficient of 0.0240 with the highest zero-llft drag coefficient slightly greater than twice the subsonic drag level. Pitchup tendencies were noted for subsonic Mach numbers at relatively high lift coefficients. Wing leading-edge droop increased the maximum lift-drag ratio approximately 8 percent at a Mach number of 0.80 but this effect was negligible at a Mach number of 0.90 and above. The configuration exhibited stable lateral characteristics over the test Mach number range.
INTRODUCTION
The present investigation is part of a general research program to evaluate the aerodynamic and hydrodynamic characteristics of water- based bomber configurations capable of flight at transonic and super- sonic speeds. Configurations of previous investigations (refs. i_ 2, and 3) had four engines with various arrangements of engine nacelles or internal ducting or both, which provided a variety of design pos- sibilities. These configurations are characterized by high fineness ratio and small frontal area and are designed in accordance with the area-rule concepts of references 4 to 6. The configuration of the present investigation had six engines and was designed for a sustained cruise Mach number of 1.8 at altitude. Four of the engines were mounted in individual nacelles on the upper surface of the wing; these nacelles were staggered to provide a smooth longitudinal area distribution. The two remaining engines were placed side by side on top of the aft section of the hull with their inlets located on top of the wing. This configuration was designed in cooperation w_th the Bureau of Aeronautics, Department of the Navy, and the aircraft industry. The hydrodynamic investigation to determine the smooth-water resist- ance, spray characteristics, and the longitudinal stability during take- offs and landings is presented in reference 7. Also included is a brief investigation of rough-water spray characteristics. In the present investigation, the stat_e longitudinal and lateral aerodynamic characteristics of the six-engine configuration were obtained over a Mach numberrange from 0.80 to 1.20. The Reynolds num- ber for the tests varied from approximately 1.65 × 106 to 1.8 × 106 based on the wing meanaerodynamic chord.
SYMBOLS
The stability system of axes is used for the presentation of the data. b wing span, in.
CD drag coefficient, Drag qS
CD,i total internal-drag coefficient of all engine nacelles based on wing area
CD,o zero-lift drag coefficient
CL lift coefficient, Lift qS
lift-curve slope per degree, averaged from C L = 0 to 0.3
CZ rolling-moment coefficient, Rolling moment qSb ct[_ rate of change of rolling-moment ,_oefficient with sideslip 8ct angle 3
Cm pitching-moment coefficient, Pitching moment about 0.25_ qS_ pitching-moment-curve slope, CmcL
Cn yawing-moment coefficient, Yawing moment q_
Cn_ rate of change of yawing-moment coefficient with sideslip $Cn angle,
Side force Cy side-force coefficient, qS
rate of change of side-force coefficient with sideslip Cyp _Cy angle,
C local wing chord, in.
wing mean aerodynamic chord, in.
i t angle of incidence of horizontal stabilizer referred to hull base line, positive when trailing edge is down, deg
(n/D)m= maximum lift-drag ratio
M free-stream Mach number q free-stream dynamic pressure_ ib/sq ft
S wing area, sq ft
£L wing chord plane angle of attack, deg
angle of sideslip, deg
DESCRIPTION OF CONFIGURATION
The general arrangement of the six-engine hull-type seaplane is shown in figure I. Pertinent characteristics and dimensions of the full-size seaplane are given in table I. General Characteristics of Full-Scale Seaplane
The six-engine configuration was desigmedwith a wing area of 1,$35 square feet and a gross weight of 220,000 pounds. Six 0renda Iroquois PS-15 turbojet engines, producing a maximumstatic thrust of 120,000 pounds without afterburners, were used.
Engine location.- Four engines were iccated in individual nacelles on top of the wing. These nacelles were staggered with the two inboard nacelles forward of the wing and the two outboard nacelles directly over the wing. The two remaining engines, which were located side by side in the after section of the hull, had their inlets located on top of the wing near the 50-percent-chord line.
Wing.- The wing had an aspect ratio of 3.0, a taper ratio of 0.333, and 0 ° sweepback of the 80-percent-chord line. A 3.5-percent-thick biconvex airfoil section, with the maximum thickness located at the 50-percent-chord line, was used. The wing incidence was 3 ° with no dihedral or twist.
Hull.- The planing bottom, with a concave-convex-shaped forebody, extended over approximately 96 percent of the fuselage length. The total length-beam ratio of the hull was 13.95 wilh a maximum be_n of 9.2 feet.
Tail group.- The horizontal tail was _ounted on top of the vertical tail. This high position was considered n_cessary for clearance of spray generated by the heavily loaded _tll.
Area Curves
The cross-sectional-area curves for tile design Mach number of 1.8 and also for a Mach number of 1.0 are pres_nted in figure 2. The area distribution for the tail group was not aw.ilable and therefore is not included in figure 2; the contribution of 1he tails to the total area curves would be relatively small. The jet mass flow subtracted from the engine nacelles is represented by an equiw_lent free-stream tube area of $0 percent of the inlet area.
Model
A O.02-scale model of the proposed si::-engine hull-type seaplane was used for this investigation. Photographs of the sting-mounted model in the Langley 8-foot transonic pressure t_Lnnel are presented in figure 3.
The model wings, tail_ hull center, aild hull aft sections were con- structed of steel; the model hull nose sec;ion and wing nacelles were made of Laminac plastic. Wing leading-edge droop was obtained by tilting a section of the leading edge of each wing semispan down8° . The drooped sections extended from 0.51b/2 to 0.94b/2 and were hinged about the 15-percent-chord line.
APPARATUSANDPROCEDURES
Tunnel
The investigation was conducted in the Langley 8-foot transonic pressure tunnel which is a single-return tunnel with a rectangular slotted test section permitting continuous operation through the tran- sonic speed range. Stagnation temperature and dewpoint controls pre- cluded the formation of condensation shocks.
Measurementsand Accuracies
Force and momentmeasurementswere obtained by a six-component electrical strain-gage balance internally mountedwithin the model. Momentsare referred to the assumedcenter of gravity (25 percent of the wing meanaerodynamic chord). All coefficients were based on the wing area of 0.734 square feet and meanaerodynamic chord of 6.438 inches. On the basis of the static calibrations of the electrical strain-gage balance and repeatability of data, the estimated accuracy of the coeffi- cients for a Mach numberof 0.80 is within the following limits:
CL ...... ±0.02 CZ ...... ±0.001 CD ...... ±0.0015 Cn ...... ±0.001 Cm ...... ±0.006 Cy ...... ±0.005
Since the accuracy is inversely proportional to the dynamic pressure_ these values decrease with Mach number. The angle of attack and angle of sideslip were determined within ±0.i °.
Internal-drag measurementswere obtained by pressure survey rakes_ consisting of total- and static-pressure tubes, located at the exit of the engine nacelles. Base-pressure measurementswere obtained from static-pressure tubes located at the base of the hull. 6
Tests
Static longitudinal tests of various model configurations were con- ducted at Mach numbers from 0.80 to 1.20 over an angle-of-attack range from approximately -4 ° to 17 ° . The configurations tested included the wing and fuselage alone and in combination with the nacelles and with both nacelles and tail; in addition, wing droop was incorporated for some tests. Most of the tests of the complete model were made with a tail incidence i t = 2° but a few were included with it = -I °. An alternate four-engine configuration was also used for a few longitudinal tests. This configuration was obtained by removing the two outboard wing nacelles of the six-engine configuration and shifting the remaining two inboard wing nacelles outboard and slightly forward. The shifted nacelles were so located that the base of each nacelle still ended at the wing trailing edge and the spanwise location was the same as that of the six-engine outboard nacelles.
Static lateral tests of the various model configurations, excluding wing droop and tail-incidence effects, were conducted over the same Mach number range with an angle-of-sideslip range from approximately -4 ° to 12 ° for an angle of attack of approximately 5.7 ° .
The stagnation pressure for all tests was maintained at 2,120 pounds per square foot. The Reynolds number based on the wing mean aerodynamic chord varied from approximately 1.65 × 106 to 1.80 × 106 .
Corrections
Boundary interference at subsonic velocities has been minimized by the slotted test section and no corrections of the type have been applied. At Mach numbers greater than 1.03 and less than 1.20, boundary reflected disturbances were present and data in this range were not taken. No cor- rections have been applied for sting-interference effects. The drag data have been adjusted to a condition of free-stream static pressure acting on the model base. The internal-drag coefficients presented in figure 4 have been removed from the drag coefficients presented in this paper.
RESULTS AND DISCUSSIDN
Longitudinal Characteristics
Lift.- The variation of lift coefficient with angle of attack for the complete six-engine configuration, as shown in the basic longitudinal 7
data presented in figures 5 to 7, exhibited nonlinearity around C L = 0 (_ _ i°). This nonlinearity of C L with _ is probably due to the flow phenomenon which occurred in the vicinity of the sharp leading edge of the wing at subsonic Mach numbers. (See ref. 8.) The addition of wing leading-edge droop tended to eliminate this nonlinearity. (See fig. 7.)
The variation of lift-curve slope CL_ over the test Mach number range for the six-engine configuration is presented in figure 8. This variation follows the usual trend through the transonic speed range.
Drag.- The variation of zero-lift drag coefficient with Mach number for the complete six-engine configuration is presented in figure 9. A value of 0.0240 for zero-lift drag coefficient was measured at subsonic Mach numbers. The highest zero-lift drag coefficient, measured at super- sonic speeds, was slightly greater than twice the value noted at subsonic speeds.
The variation of maximum lift-drag ratio with Mach number for the six-engine configuration with and without wing leading-edge droop is presented in figure i0. The six-engine configuration without wing leading-edge droop had a maximum lift-drag ratio of approximately 7.5 at subsonic speeds and followed the usual trend of decreasing with Mach nu_er. The addition of wing leading-edge droop increased the maximum lift-drag ratio approximately 8 percent at a Mach number of 0.$0; however, this beneficial effect was negligible at a Mach number of 0.90 and above.
Pitchin_ moment.- The complete six-engine configuration experienced pitchup tendencies (fig. 7) at lift coefficients above 0.7 up to a Mach n_r_er of 0.93. The wing leading-edge droop slightly increased the lift coefficient at which pitchup occurred.
The pitching-moment-curve slope Cmc L for the complete six-engine configuration is presented in figure ii. The slopes were determined near a trim-lift coefficient of approximately 0.3. The rearward movement of the aerodyn_aic center is noted to be approximately 15 percent of the wing mean aerodyn_nic chord for the test Mach nu_._ber range.
Alternate four-engine configuration.- Results of tests of the alternate four-engine configuration are presented in figures 12 and 13. The basic longitudinal data of figure 12 show the same nonlinear varia- tion of lift coefficient with angle of attack near zero lift that was noted for the six-engine configuration (figs. 5 to 7). The variation of pitching-moment coefficient over the test lift-coefficient range revealed no pitchup tendencies for the alternate four-engine configuration. The variation of the longitudinal aerodynamic characteristics with Mach num- _r, presented in figure 13, show the usual trends through the transonic speed range. Lateral Characteristics
The lateral stability characteristics of the six-engine configura- tion over the test Mach numberrange at _ _ 5.7° are presented in fig- ure 14. The lateral stability derivatives are presented in figure 15.
In general, the complete six-engine configuration exhibited stable characteristics over the test Machnur_berrange. However_the addition of the wing nacelles had a slight unstable contribution to the confib_ration.
CONCLUDINGRE_RKS
The transonic aerodynamic characteristics of a six-engine hull-type seaplane configuration designed for supersonic flight have been investi- gated. The results of the tests show tl_t: i. The configuration had a subsonic zero-lift drag coefficient of 0.0=_0_O I with the highest zero-lift drag coefficient measured being slightly greater than twice the subsonic level.
2. Pitchup tendencies were noted for subsonic Mach numbers at relatively high lift coefficients.
_. Wing leading-edge droop increased the maximum lift-drag ratio approxin_tely $ percent at a Mach n_r_er of 0.$0 but this effect was negligible at a Mach nu_._er of 0.90 and above.
4. The configuration exhibited stable lateral characteristics over the test _4ach n_nber range.
Langley Research Center, National Aeronautics and Space A_ninistration, Langley Field, Va., December i, 1959. 9
REFERENCES
i. Petynia, William W., Hasson, Dennis F., and Spooner, Stanley H.: Aerodynamic and Hydrodynamic Characteristics of a Proposed Super- sonic MultiJet Water-Based Hydro-Ski Aircraft With a Variable- Incidence Wing. NACA RML57G05, 1957.
2. Bielat, Ralph P._ Coffee, Claude W., Jr._ and Petynia_ William W.: Aerodynamic and Hydrodynamic Characteristics of a Deck-lnlet _iti- jet Water-Based-Aircraft Configuration Designed for Supersonic Flight. NACA RML56HOI, 1956.
3. Olson, Roland E., and Bielat, Ralph P.: An Aerodynamic and Hydro- dynamic Investigation of Two MultiJet Water-Based Aircraft Having Low Transonic Drag Rise. NACA RML55AIIa, 1955.
4. Whitcomb, Richard T.: A Study of the Zero-Lift Drag-Rise Character- istics of Wing-Body Combinations Near the Speed of Sound. NACA Rep. 1273, 1956. (Supersedes NACA RM L52H08.)
5. Whitcomb, Richard T._ and Fischetti, Thomas L.: Development of a Supersonic Area Rule and an Application to the Design of a Wing- Body Combination Having High Lift-to-Drag Ratios. NACA RM L53H31a, 1953.
6. Whitcomb, Richard T.: Some Considerations Regarding the Application of the Supersonic Area Rule to the Design of Airplane Fuselages. NACA RM L56E23a, 1956.
7. Coffee, Claude W., Jr.: Hydrodynamic Characteristics of a Model of a Proposed Six-Engine Hull-Type Seaplane Designed for Supersonic Flight. NACA RM L58EI3, 1958.
8. Lindsey, W. F., Daleyj Bernard N., and Humphreys, Milton D.: The Flow and Force Characteristics of Supersonic Airfoils at High Subsonic Speeds. NACA TN 1211, 1947. i0
TABLE I.- PERTINENT cHARACTERISTICS _ND DIMENSIONS
OF THE FUIT_SIZE SIX-ENGINE S _APLANE
General: Gross weight, ib ...... 220,000 Wing area, sq ft ...... 1,839 Engines, Orenda Iroquois PS-I} ...... 6 Takeoff t_ust (wlth<_ut afterburners), ib ...... 120,OOO Wing loading, ib/sq ft ...... 120 Takeoff thrust-welght ratio ...... 0.545
Wing: Span, ft ...... 74.2 Airfoil section ...... Biconvex Thickness, percent chord ...... 9.5 Aspect ratio ...... 5.O Taper ratio ...... 0.353 Sweepback (0.25c), deg ...... 20.17 Sweepback (O.8Oe), deg ...... O Incidence (wing root to base llne), deg ...... 3 Geometric twist, deg ...... O Dihedral, deg ...... O Root chord, ft ...... 37.16 Mean aerodynamic chord, ft ...... 26._2 Tip chord, ft ...... 12.53
Hull: Length, overall, ft ...... 134.09 Beam at chines, maximum ft ...... 9.2 Height, maximum including hull nacellesj ft ...... 15
Horizontal tail: Span, ft ...... 52.7 Airfoil section ...... Biconvex Thickness, percent chord ...... 3.5 Area, sq ft ...... 555 Aspect ratio ...... 3.0 Taper ratio ...... 0.4 Sweepback (0.25c), deg ...... 17.3 Sweepback (0.80c), deg ...... 0 Dihedral, deg ...... 15 Tail length (from 0.25 wing _ to 0.25 tail 5 parall_l to wing root chord), ft ...... 65.08
Vertical tail: Airfoil section ...... NACA 64A006 Area, sq ft ...... 242.6 Aspect ratio ...... i.i Sweepback (0.25c), deg ...... 50
Area curves: Maeh number, 1.0 Maximum cross-sectional area, sq ft ...... 157.9 Maximum diameter of equivalent body, ft ...... 14.2 Length, ft ...... 134.1 Fineness ratio of equivalent body ...... 9.5 Mach number, 1.8 Maximum cross-sectional area, sq ft ...... 135.0 Maximum diameter of equivalent body, ft ...... 13.i Length, ft ...... 134.1 Fineness ratio of equivalent body ...... 10.2 Total surface area, sq ft ...... 9,980 1Z
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Figure 5.- Longitudinal aerodynamic characteristics of the six-engine configurations. 16
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Figure _.- Continued. 17
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Figure 6.- Concluded. 23
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Figure 12,- Longitudinal aerodynamic charact(ristics of the four-engine configurations. 29
(b) M = 0.90 and 0.95.
Figure 12.- Continued. 30
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Figure 15.- Variation with Mach number of the lateral stability deriva- tives for the six-engine configurations. _ _ 5.7 °.
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