Showing papers on "Drag divergence Mach number published in 1977"

Application of numerical optimization to the design of supercritical airfoils without drag-creep

Abstract: Recent applications of numerical optimization to the design of advanced airfoils for transonic aircraft have shown that low-drag sections can be developed for a given design Mach number without an accompanying drag increase at lower Mach numbers. This is achieved by imposing a constraint on the drag coefficient at an off-design Mach number while the drag at the design Mach number is the objective function. Such a procedure doubles the computation time over that for single design-point problems, but the final result is worth the increased cost of computation. The ability to treat such multiple design-point problems by numerical optimization has been enhanced by the development of improved airfoil shape functions. Such functions permit a considerable increase in the range of profiles attainable during the optimization process.

An overview of concepts for aircraft drag reductions

Abstract: A current overview of aerodynamic drag reduction concepts which have potential for reducing aircraft fuel consumption is presented. The discussion shows where the greatest percentages of aircraft fuel is burned and what areas have the greatest potential for fuel conservation. The paper deals with aerodynamic improvements and touches only briefly on structural and propulsion improvements. Concepts for reducing pressure drag (i.e., roughness, wave, interference, and separation drag), drag due to lift/induced drag, and skin-friction drag at subsonic and supersonic speeds are emphasized.

Two-dimensional aerodynamic characteristics of several rotorcraft airfoils at Mach numbers from 0.35 to 0.90

Abstract: An investigation was conducted in the Langley 6- by 28-inch transonic tunnel and the 6- by 19-inch transonic tunnel to determine the two-dimensional aerodynamic characteristics of several rotorcraft airfoils at Mach numbers from 0.35 to 0.90. The airfoils differed in thickness, thickness distribution, and camber. The FX69-H-098, the BHC-540, and the NACA 0012 airfoils were investigated in the 6- by 28-inch tunnel at Reynolds numbers (based on chord) from about 4.7 to 9.3 million at the lowest and highest test Mach numbers respectively. The FX69-H-098, the NLR-1, the BHC-540, and the NACA 23012 airfoils were investigated in the 6- by 19-inch tunnel at Reynolds numbers from about 0.9 to 2.2 million at the lowest and highest test Mach numbers respectively.

TRANDES: A FORTRAN program for transonic airfoil analysis or design

Abstract: A program called TRANDES is presented that is used for the analysis of steady, irrotational transonic flow over specified two-dimensional airfoils in free air or for the design of airfoils having a prescribed pressure distribution, including the effects of weak viscous interaction. Instructions on program usage, listings of the program, and sample cases are given.

Constrained Flow Past Cavitating Bluff Bodies

Abstract: An experimental program has been carried out to determine the effects of wall interference on the drag and vortex shedding characteristics of cavitating two-dimensional triangular prisms and circular cylinders. The former shapes were chosen to eliminate effects of Reynolds number in interpreting the results. Direct pressure measurements were made to estimate the drag force. The vortex shedding frequency of the cavitating bodies was recorded with the help of a pressure transducer. The gap velocity u1 and the jet contraction velocity uj are shown to be the proper velocity scales to form the drag coefficients and Strouhal numbers for the bluff shapes tested. The drag coefficient was found to increase due to wall interference effects when partial cavitation conditions prevailed. The trend of the drag coefficient data indicated that wall interference effects are relatively small at very low cavitation numbers (σ → σch ) which correspond to choking conditions. As choking conditions are reached, the vortex shedding from the cavitating source becomes intermittent and finally vortex shedding ceases.

Form drag, skin friction, and vortex shedding frequencies for subsonic and transonic crossflows on circular cylinder

Abstract: A series of wind-tunnel tests covering a range of Mach numbers and Reynolds numbers in subsonic and transonic flows was conducted on a circular cylinder placed normal to the flow. Form drag coefficients were determined from surface-pressure measurements and displayed as a function of Mach number to show the drag rise phenomenon. Buried wire gages arranged on the model surface were used to measure skin-friction distributions and vortex-shedding frequencies at different flow conditions. It was found that detectable periodic shedding ceases above M = 0.9. The measured skin-friction distributions indicate the positions of mean separation points clearly; these values are documented for the different flow conditions.

Wind-tunnel pressure data at Mach numbers from 1.6 to 4.63 for a series of bodies of revolution at angles of attack from -4 deg to 60 deg

Abstract: The tabulated results of wind tunnel pressure tests are presented without analysis. The data were obtained for a series of six bodies of revolution at Mach numbers of 1.6, 2.3, 2.96, and 4.63 for angles of attack from -4 deg. to 60 deg. The Reynolds number used for these tests was 6.6 x 6/million per meter.

Methods for Reducing Subsonic Drag Due to Lift

Abstract: The results of repeat experimental research on methods for reducing subsonic drag due to lift are discussed. The NASA supercritical airfoils and their application to structurally practical wings with increased aspect radio are described. A design approach and experimental results for wing-tip-mounted winglets are presented. Several methods for utilizing the thrust of jet engines to provide reductions in the drag due to lift are also discussed.

Mean velocity and turbulent energy closures for flows with drag reduction

Abstract: Earlier models for flows with drag reduction are reviewed and compared with a new energy‐dissipation closure which describes the observed effects of drag reduction on the mean velocity, turbulent energy, and turbulent length‐scale distributions, and suggests a new maximum drag reduction law related to the onset of drag reduction.

Supersonic-Transonic Flow Generated by a Thin Airfoil in a Stratified Atmosphere

Abstract: This paper deals with the steady, frictionless, nonheatconducting flow field of a thin airfoil moving supersonically in an atmosphere with a weak wind gradient. Because of this wind gradient, the flow gradually becomes transonic in the far field below the airfoil. It is shown that a solution which describes both supersonic and transonic regimes can be derived by matching two expansions. One of these describes the flow in the “outer” supersonic region, and is a multiple-scale expansion which can be calculated analytically. Since this expansion becomes singular as the local Mach number approaches unity, one then considers an “inner” transonic expansion valid in a layer of the atmosphere which is thin compared to the distance between the airfoil and the sonic line. Although the solution of the transonic equation cannot be calculated analytically for the present case where there are shocks in the flow, an asymptotic representation of this solution is derived in the supersonic region of the far-field, and this...

Calculations of Some Unsteady Transonic Flows About the NACA 64A006 and 64A010 Airfoils

Abstract: : Unsteady transonic flows over airfoils were calculated using a program based on the unsteady Euler equations. The approximate numerical solutions were obtained using an explicit (Lax-Wendroff) difference scheme. Calculations were made for the 64A006 airfoil at zero angle-of-attack in flows with Mach number 0.822, 0.854 and 0.875. Three unsteady flows caused by a quarter-chord flap oscillating with an amplitude on the order of one degree at specific reduced frequencies (k = omega C/U sub infinity) near 0.4 were analyzed. The results were compared with other available calculations and the experimental data of Tijdeman. Responses for an airfoil in a ventilated-wall wind tunnel rather than in a free-stream were also calculated. Exploratory study of the changes to be expected with a weakened shock (an attempt to simulate a main effect of the interaction of the shock with airfoil boundary layer) was done for one case at Mach number 0.875. Oscillatory flows over the 64A010 airfoil at zero angle-of-attack in a Mach 0.80 stream were also calculated. Plunging motion at reduced frequency 0.4 and pitching motions at reduced frequencies 0.4 and 0.5 were calculated.

Flight-measured lift and drag characteristics of a large, flexible, high supersonic cruise airplane

Abstract: Flight measurements of lift, drag, and angle of attack were obtained for the XB-70 airplane, a large, flexible, high supersonic cruise airplane. This airplane had a length of over 57 meters, a takeoff gross mass of over 226,800 kilograms, and a design cruise speed of Mach 3 at an altitude of 21,340 meters. The performance measurements were made at Mach numbers from 0.72 to 3.07 and altitudes from approximately 7620 meters to 21,340 meters. The measurements were made to provide data for evaluating the techniques presently being used to design and predict the performance of aircraft in this category. Such performance characteristics as drag polars, lift-curve slopes, and maximum lift-to-drag ratios were derived from the flight data. The base drag of the airplane, changes in airplane drag with changes in engine power setting at transonic speeds, and the magnitude of the drag components of the propulsion system are also discussed.

Effect of nonlifting empennage surfaces on single-engine afterbody/nozzle drag at Mach numbers from 0.5 to 2.2

Abstract: The effect of empennage interference on the drag characteristics of a model with a single engine fighter aft end with convergent-divergent nozzles was studied. The dry and maximum afterburning nozzle power settings were investigated. A high pressure air system was used to provide jet total pressure ratios up to 20.0. In an attempt to quantify and reduce adverse empennage interference and decrease aft-end drag, several empennage arrangements (variable tail surface location), contour bump configurations, and locally contoured afterbodies were investigated. The results of the investigation indicate that empennage interference effects can be significant at transonic and supersonic speeds. The most effective means of reducing adverse empennage interference is the proper relocation of individual tail surfaces. The aft or conventional empennage arrangement produced the highest aft-end drag at all conditions investigated.

Theoretical Aspects of Dromedaryfoil.

Abstract: : A new airfoil design (called a dromedaryfoil) has been developed using a single hump on a modified supercritical airfoil for limiting the center of pressure excursion and maximizing the drag divergence Mach number. Derivation of the hump is based on isentropic compression in the fore part and incipient separation in the rear. The former leads to a weakened shock wave and the latter to high pressure recovery after the shock. The shock will theoretically locate at the peak of the hump to form a fixed pressure pattern under different flight speeds. The shock foot will be inclined at a deflection angle of the hump measured from the normal of the fore hump surface at the peak. Theoretical results indicate considerably shorter center-of-pressure travel for a dromedaryfoil than for a supercritical airfoil with equal wave drag. However, improper humping would be penalized by increased wave drag. At high supercritical flows, the shock strength would be limited by (M sub 1 sin beta)max = 1.483. Experimental verification of theoretical predictions is planned. (Author)

Computational Research on Inviscid, Unsteady, Transonic Flow Over Airfoils.

Abstract: : The inviscid transonic flow over an NACA 64A410 airfoil oscillating in pitch in a Mach 072 stream was calculated with a program based on the unsteady Euler equations The airfoil oscillates about a mid-chord axis with attitude alpha = 1 deg + or - 1 deg at reduced frequency k = omega C/U(infinty) = 02 The effects of two approximations made in the analysis, handling of boundary conditions at the airfoil surface and at the perimeter of the computation field, have been studied (Author)

Aerodynamic Characteristics at Mach Numbers from 0.6 to 2.16 of a Supersonic Cruise Fighter Configuration with a Design Mach Number of 1.8

Abstract: An investigation was made in the Langley 8-foot transonic tunnel and the Langley Unitary Plan wind tunnel, over a Mach number range of 0.6 to 2.16, to determine the static longitudinal and lateral aerodynamic characteristics of a model of a supersonic-cruise fighter. The configuration, which is designed for efficient cruise at Mach number 1.8, is a twin-engine tailless arrow-wing concept with a single rectangular inlet beneath the fuselage and outboard vertical tails and ventral fins. It had untrimmed values of lift-drage ratio ranging from 10 at subsonic speeds to 6.4 at the design Mach number. The configuration was statically stable both longitudinally and laterally.

Nacelle Integration Study on a Mach-2.2 Supersonic Cruise Aircraft

Abstract: Results of a propulsion system integration study performed on a Mach-2.2 advanced supersonic cruise aircraft are discussed. A study configuration developed in a joint NASA-Douglas supersonic technology program was used as the baseline airframe to study the detailed problems of inlet-nacelle -airframe integration at Mach 2.2. Numerous inlet-nacelle combinations were examined in a preliminary screening study. Promising configurations were evaluated in a nacelle installation study in which structural weight and installed wave drag were traded leading to the selection of an axisymmetric single-engine pod installation as the most promising configuration. A detailed nacelle shape study was conducted, and a wing reflex was designed. A cooperative NASA-Douglas wind-tunnel test of the refined nacelle with both mixed and external compression inlets was conducted with the nacelles installed on both a refined baseline wing and a reflexed wing. An installation drag penalty equal to 4.3% of the baseline wing-body drag was observed for the external compression inlet over the mixed compression inlet. Wing reflexing improved the trimmed wing-body-nacelle drag by 3.0% of the wing-body drag. Good agreement was observed between calculated and experimental increments in induced drag due to nacelle installation.

A Comparison Between a Simple Drag Formula and Experimental Drag Data for Bodies of Revolution

Abstract: : A simplified drag formula to predict the drag of a small number of axisymmetric forms is systematically applied to seven series of model forms comprising nearly fifty bodies. Calculations of form (or residual) drag are compared to available experimental data in order to determine the usefulness of the method for predictive purposes. The formula is shown to exhibit very little sensitivity to changes in most body parameters such as the length-of-stern to diameter ratio, nose and tail radii parameters, and the prismatic coefficient. For some parameters, such as length to diameter ratio, the length of bow section to diameter ratio, however, the formula is sometimes able to discriminate bodies having high values of form drag. It is concluded that the simple drag formula may not be reliably used for estimating the relative form drags of bodies of revolution.

Experimental evaluation of nacelle-airframe interference forces and pressures at Mach numbers of 0.9 to 1.4

Abstract: Detailed interference force-and-pressure data were obtained on a representative supersonic transport wing-body-nacelle combination at Mach numbers of 0.9 to 1.4. The basic model consisted of a delta wing-body aerodynamic model with a length of 158.0 cm (62.2 in.) and a wingspan of 103.6 cm (40.8 in.) and four independently supported nacelles positioned beneath the model. The experimental program was conducted in the Ames 11- by 11-Foot Wind Tunnel at a constant unit Reynolds number. The primary variables examined included Mach number, angle of attack, nacelle position, and nacelle mass-flow ratio. Under the most favorable conditions, the net interference drag was equal to 50 percent the drag of four isolated nacelles at M = 1.4, 75 percent at M = 1.15, and 144 percent at M = 0.90. The overall interference effects were found to be rather constant over the operating angle-of-attack range of the configuration. The effects of mass-flow ratio on the interference pressure distributions were limited to the lip region of the nacelle and the local wing surface in the immediate vicinity of the nacelle lip. The net change in the measured interference forces resulting from variations in the nacelle mass-flow ratio were found to be quite small.

The transonic multi-foil Augmentor-Wing

Abstract: The paper describes the development of a transonic blown multi-foil Augmentor-Wing airfoil section that has a thickness/chord (t/c) value of 0.18. In comparison with an unblown single-foil supercritical section of the same overall t/c the new multi-foil section is characterized by an increased drag rise Mach number, increased buffet boundaries, and a reduction in 'effective' drag due to blowing. Potential advantages of the Augmentor-Wing are considered and the testing of three high-speed models in a trisonic pressurized wind tunnel (possessing a two-dimensional transonic insert) is discussed. The data indicate that a very thick wing is feasible since separations toward the rear of the main foil can be controlled both by shroud location and augmentor blowing.

Subsonic and supersonic aerodynamic characteristics of a supersonic cruise fighter model with a twisted and cambered wing with 74 deg sweep

Abstract: A wind tunnel investigation has been conducted to determine the longitudinal and lateral aerodynamic characteristics of a model of a supersonic cruise fighter configuration with a design Mach number of 2.60. The configuration is characterized by a highly swept arrow wing twisted and cambered to minimize supersonic drag due to lift, twin wing mounted vertical tails, and an aft mounted integral underslung duel-engine pod. The investigation also included tests of the configuration with larger outboard vertical tails and with small nose strakes.

Supersonic Wave Drag for Nonplanar Singularity Distributions

Abstract: Introduction T paper derives a planar counterpart to von Karman's * well-known drag formula for axisymrnetric lineal source distributions but which includes the effect of doublets. It provides an explicit evaluation of the author's formal results (for general nonplanar singularity distributions) in the planar case. Current wave drag prediction methods in supersonic aerodynamics are based on the classical analyses of von K&rman and Hayes. An exact result due to Hayes applies only to the source problem and is summarized easily. For sources of density/(Q), where Q is the source coordinate, define an equivalent density/such that

Transonic wind-tunnel investigation of the maneuver potential of the NASA supercritical wing concept, phase 1

Abstract: An investigation was conducted in the NASA Langley 8-foot transonic pressure tunnel at Mach numbers from 060 to 0975 with a variable-wing-sweep airplane model in order to evaluate a series of wings designed to demonstrate the maneuver potential of the supercritical airfoil concept Both conventional and supercritical wing designs for several planform configurations were investigated with wing sweep angles from 160 deg to 725 deg, depending on Mach number and wing configuration The supercritical wing configuration showed significant improvements over the conventional configurations in drag-divergence Mach number and in drag level at transonic maneuver conditions

An investigation of induced drag reduction through over-the-wing blowing

Abstract: An investigation has been conducted in the Langley 16-foot transonic tunnel to verify analytically predicted benefits in climb and cruise performance due to blowing the jet exhaust over the wing for a transport configuration. A wing-body model - powered-nacelle rig combination was tested at Mach numbers of 0.5 and 0.8 at angles of attack from -2 to 4 deg and jet total pressure ratios from jet off to 3 or 4 (depending on Mach number) for a variety of nacelle locations relative to the wing. Results from this investigation show that the induced drag for the wing-body (nacelles were nonmetric) was reduced for virtually all configurations. In addition to the experimental results, comparisons of the data with available prediction methods are included to show their validity and capabilities.

Drag evaluation of the Bellanca Skyrocket II

Abstract: The Bellanca Skyrocket II, possessor of five world speed records, is a single engine aircraft with high performance that has been attributed to a laminar flow airfoil and an all composite structure. Utilization of composite materials in the Skyrocket II is unique since this selection was made to increase the aerodynamic efficiency of the aircraft. Flight tests are in progress to measure the overall aircraft drag and the wing section drag for comparison with the predicted performance of the Skyrocket. Initial results show the zero lift drag is indeed low, equalling 0.016.

Comment on 'Drag coefficient of spheres in continuum and rarefied flows'

Abstract: A paper by Henderson (1976) provides a method of predicting experimental sphere drag data. This approach uses two equations for the drag coefficient, one for relative Mach number less than one, one for relative Mach number greater than 1.75. For relative Mach numbers between these limits a linear interpolation procedure is followed. In a comment on this paper, it is claimed, on the basis of comparing predictions with experimental results, that a method proposed by Walsh (1975) gives better predictions of the drag coefficient for relative Mach numbers less than 1.75, provided that a modification of the procedure is made for relative Mach numbers less than 0.1. For values over 1.75, both methods are considered equally accurate. In a reply to this comment, it is agreed that the Walsh method is more accurate when Reynolds numbers are within a range between 20 and 200, and Mach numbers are between 0.5 and 1.25. Presumed errors and possible limitations in the Walsh procedure for predicting drag coefficients are discussed.

Effect of an integrated scramjet installation on the subsonic performance of an aircraft designed to Mach 6 cruise

Abstract: Various factors contributing to the high drag caused by the installation of a six-module scramjet engine were determined from wind tunnel tests at Mach numbers from 0.2 to 0.7. Methods for alleviating this drag were also explored. The external exhaust nozzle, required for good cruise performance, was a major contributor. Of the drag produced by the engine modules, a significant fraction was attributable to wall divergence in the combustor. Good drag simulation could be achieved by using a single fuel injection strut having approximately the same cross-sectional area as the three used on the full-scale engine. External exhaust nozzle fences had a small but beneficial effect on maximum L/D and a flap which diverted the flow away from the inlet was effective in decreasing drag but only at low angles of attack.

An Investigaion of Boattail Geometry and Reynolds Number Effects on Forebody and Afterbody Drag at Transonic Mach Numbers

Abstract: : An experimental investigation was conducted in the AEDC 16-foot Transonic Wind Tunnel (16T) to determine both Reynolds number and nozzle afterbody configuration effects on model forebody and afterbody drag. The model was a sting-mounted body of revolution with interchangeable contoured, cylindrical, and 15-deg boattail configurations. Pressure and force data were obtained at Mach numbers from 0.60 to 1.40 and at unit Reynolds numbers from 1, 470,000 per foot to 5,300,000 per foot. The experimental results showed that large variations in afterbody drag levels produced no significant change in forebody drag. The data also revealed that all three configurations exhibited little or no Reynolds number dependence subsonically and that only the 15-deg boattail afterbody was affected by Reynolds number supersonically. It was also demonstrated that data precision and wind tunnel calibration can have a significant effect on model drag and should be given careful consideration.

Optimal Projectile Shapes for Minimum Total Drag

Abstract: : An analytical method for theoretically predicting the projectile shape which yields the minimum total drag for a fixed length, diameter, and supersonic Mach number (2< or = free stream mach no < or = 6) is derived. The pressure drag was estimated by modified Newtonian theory on the nose and Prandtl-Meyer expansion on the afterbody. The skin-friction drag was approximately by Van Driest method and the base drag by a semiempirical technique. The drag on the forebody is optimized using a new numerical technique and on the afterbody by using the method of steepest descent. The optimum body shape has a base diameter of about 70 percent of the maximum diameter and a forebody length varying between 60 and 80 percent percent of the total length depending on the Mach number and overall fineness ratio. The forebody ogive shape lies between the well-known hypersonic optimum 2/3- and 3/4-power law ogives and the afterbody is conical. Results further indicate that a change of 5 percent in nose length from the optimum results in only about a 1-percent drag penalty.