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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1944

The "limiting line" in mixed subsonic and supersonic flow of compressible fluids

It is well known that the vorticity for any fluid element is constant if the fluid is non-viscous and the change of state of the fluid is isentropic. When a solid body is placed in a uniform stream, the flow far ahead of the body is irrotational. Then if the flow is further assumed to be isentropic, the vorticity will be zero over the whole filed of flow. In other words, the flow is irrotational. For such flow over a solid body, it is shown by Theodorsen that the solid body experiences no resistance. If the fluid has a small viscosity, its effect will be limited in the boundary layer over the solid body and the body will have a drag due to the skin friction. This type of essentially isentropic irrotational flow is generally observed for a streamlined body placed in a uniform stream, if the velocity of the stream is kept below the so-called "critical speed." At the critical speed or rather at a certain value of the ratio of the velocity of the undisturbed flow and the corresponding velocity of sound, shock waves appear. This phenomenon is called the "compressibility bubble." Along a shock wave, the change of state of the fluid is no longer isentropic, although still adiabatic. This results in an increase in entropy of the fluid and generally introduces vorticity in an originally irrotational flow. The increase in entropy of the fluid is, of course, the consequence of changing part of the mechanical energy into heat energy. In other words, the part of fluid affected by the shock wave has a reduced mechanical energy. Therefore, with the appearance of shock waves, the wake of the streamline body is very much widened, and the drag increases drastically. Furthermore, the accompanying change in the pressure distribution over the body changes the aerodynamic moment acting on it and in the case of an airfoil decreases the lift force. All these consequences of the breakdown of isentropic irrotational flow are generally undesirable in applied aerodynamics. Its occurrence should be delayed as much as possible by modifying the shape or contour of the body. However, such endeavor will be very much facilitated if the cause or the criterion for the breakdown can be found first.

MACH NUMBER↗
NASA NTRS1990

Supersonic aircraft drag reduction

This paper reviews aerodynamic drag reduction for friction, wave and vortex drag associated with supersonic cruise aircraft and suggests approaches and research directions. Suction laminar flow control may also enable improved low-speed, high-lift systems, improved lift-to-drag ratio for subsonic cruise, reduced parasitic viscous drag for favorable interference wave drag reduction approaches, and turbulent skin friction reduction via slot injection. Flow separation control at cruise proffers opportunities for increased leading-edge thrust, increased lift increment from upper surface, increased fuselage lift/camber for wave drag-due-to-lift reduction, improved performance of various favorable interference wave drag reduction schemes, as well as possibly better low-speed, high-lift systems and wing cruise performance.

Bushnell, D. M.↗
NASA NTRS1973

Review of drag measurements from flight tests of manned aircraft with comparisons to wind-tunnel predictions

In-flight studies of the overall and local components of drag of many types of aircraft were conducted. The primary goal of these studies was to evaluate wind-tunnel and semiempirical prediction methods. Some evaluations are presented in this paper which may be summarized by the following observations: Wind-tunnel predictions of overall vehicle drag can be accurately extrapolated to flight Reynolds numbers, provided that the base drag is removed and the boattail areas on the vehicle are small. The addition of ablated roughness to lifting body configurations causes larger losses in performance and stability than would be expected from the added friction drag due to the roughness. Successful measurements of skin friction have been made in flight to Mach numbers above 4. A reliable inflatable deceleration device was demonstrated in flight which effectively stabilizes and decelerates a lifting aircraft at supersonic speeds.

Pyle, J. S.↗
NASA NTRS1984

Effects of tail span and empennage arrangement on drag of a typical single-engine fighter aft end

An investigation was conducted in the Langley 16 foot Transonic Tunnel to determine the effects of tail span and empennage arrangement on drag of a single engine nozzle/afterbody model. Tests were conducted at Mach numbers from 0.50 to 1.20, nozzle pressures frm 1.0 (jet off) to 8.0, and angles of attack from -3 to 9 deg, depending upon Mach numbers. Three empennage arrangements (aft, staggered, and forward) were investigated with several different tail spans. The results of the investigation indicate that tail span and position have a significant effect on the drag at transonic speeds. Unfavorable tail interference was largely due to the outer portion of the tail surfaces. The inner portion near the nozzle and afterbody did little to increase drag other than surface skin friction. Tail positions forward of the nozzle generally had lower tail interference.

Burley, J. R., II↗
NASA NTRS1992

Aircraft Drag Reduction

This paper reviews aerodynamic drag reduction for friction, wave and vortex drag associated with supersonic cruise aircraft and suggests approaches and research directions which cover the spectrum from possibly significant to revolutionary. Various synergisms are also included, especially in regard to suction laminar flow control and flow separation control at cruise. The former may also enable improved low speed, high lift systems, lift to drag ratio for subsonic cruise, reduced parasitic viscous drag for favorable interference wave drag reduction approaches, and turbulent skin friction reduction via slot injection. Flow separation control at cruise proffers opportunities for increased leading edge thrust, increased lift increment from upper surface, increased fuselage lift/camber for wave drag due to lift reduction, improved performance of various favorable interference wave drag reduction schemes, as well as possibly better low speed, high-schemes and wing cruise performance. More speculative approaches include multi-stage aircraft, thrust vectoring for lift augmentation, as well as trim, and a number of conceptual wave and vortex drag reduction schemes.

D M Bushnell↗
NASA NTRS1978

Drag and heat transfer on surfaces with small longitudinal fins

An investigation is conducted regarding the drag characteristics of longitudinally ribbed surfaces using various configurations which attempt to confine the turbulent wall bursts to regions of small transverse extent in the initial birth region. The ratio of heat transfer increase to drag increase for the high area ratio rib configurations is determined and the heat transfer efficiency of the present surfaces is compared with data found in the literature. The drag data obtained indicate local 'wetted area skin friction' reductions but due to the increased surface area the planform drag is increased in most instances. There is no indication of drag reduction with the rectangular ribbed models that had sharp leading and fin tip edges. Heat transfer data indicate that several of the triangular grooved models have efficiencies approximately 10 percent higher than that of a smooth flat plate.

Walsh, M. J.↗
NASA NTRS2000

Review of Skin Friction Measurements Including Recent High-Reynolds Number Results from NASA Langley NTF

This paper reviews flat plate skin friction data from early correlations of drag on plates in water to measurements in the cryogenic environment of The NASA Langley National Transonic Facility (NTF) in late 1996. The flat plate (zero pressure gradient with negligible surface curvature) incompressible skin friction at high Reynolds numbers is emphasized in this paper, due to its importance in assessing the accuracy of measurements, and as being important to the aerodynamics of large scale vehicles. A correlation of zero pressure gradient skin friction data minimizing extraneous effects between tests is often used as the first step in the calculation of skin friction in complex flows. Early data compiled by Schoenherr for a range of momentum thickness Reynolds numbers, R(sub Theta) from 860 to 370,000 contained large scatter, but has proved surprisingly accurate in its correlated form. Subsequent measurements in wind tunnels under more carefully controlled conditions have provided inputs to this database, usually to a maximum R(sub Theta) of about 40,000. Data on a large axisymmetric model in the NASA Langley National Transonic Facility extends the upper limit in incompressible R(sub Theta) to 619,800 using the van Driest transformation. Previous data, test techniques, and error sources ar discussed, and the NTF data will be discussed in detail. The NTF Preston tube and Clauser inferred data accuracy is estimated to be within -2 percent of a power-law curve fit, and falls above the Spalding theory by 1 percent at R(sub Theta) of about 600,000.

Watson, Ralph D.↗
NASA NTRS1984

Performance of large-eddy breakup devices at post-transitional Reynolds numbers

Large-eddy alteration techniques were used to modify the drag characteristics of the turbulent boundary layer. Local skin-friction coefficients were measured for single and multi-element arrays of thin plates suspended in a fully turbulent, flat plate, boundary layer at R(theta) = 3,000. Effects of these devices were measured more than 300 boundary-layer thicknesses downstream to R(theta) = 17,000. It was found that although most of the devices reduced the local skin friction immediately downstream, this reduced skin-friction region persisted for only 100-120 delta-0, after which a rapid rise in skin-friction coefficient occurred, often exceeding flat plate values. Net drag reductions were obtained only for tandem configurations and these reductions were sensitive to device spacing and height. The maximum net drag reduction reported was approximately 7 percent with a device spacing of 10 delta-0 at a height of 0.8 delta-0.

Anders, J. B.↗
NASA NTRS1990

Boundary layer manipulators at high Reynolds numbers

Airfoil large-eddy breakup (LEBU) devices were tested on an axisymmetric body in the Langley Two Tank up to speeds of 50 ft/sec. NACA-0009, NACA-2412, E-193, and Clark Y contours were tested in single and tandem configurations. At the higher Reynolds numbers local skin friction downstream of the devices showed minimal reductions O (10 percent) and total body drag was increased 1 to 3 percent. At lower Reynolds numbers skin-friction reductions as large as 25 percent were measured and total body drag tended toward net reductions. The loss of effectiveness with increasing Reynolds number of conventional, outer layer devices suggests a decoupling of the outer and inner scales in high Reynolds number turbulent boundary layers.

Anders, J. B.↗
NASA NTRS1953

Recent Results Pertaining to the Application of the "Area Rule"

This paper is concerned primarily with the application of the "area rule" to the interpretation and improvement of the drag-rise characteristics of wing-body combinations at transonic and moderate supersonic speeds. Consideration of the general physical nature of the flow at transonic speeds, together with comparisons of the flow fields and drag-rise characteristics for wing-body combinations and bodies of revolution has led to the conclusion that near the speed of sound the drag rise for a thin low-aspect-ratio wing-body combination is primarily dependent on the axial distribution of cross-sectional area normal to the airstream (ref. 1). (The drag rise, sometimes referred to as pressure drag, is the difference between the drag level near the speed of sound and the drag level at subsonic speeds where the drag is due primarily to skin friction.) In order to illustrate the concept, figure 1 shows a wing-body combination and a body of revolution. A typical cross section normal to the airstream for the wing-body combination is shown at AA. The cross-sectional area of the wing is wrapped around the body of revolution so that the body has the same cross-sectional area at BB. All the other cross-sectional areas of the body of revolution are the same as those for the wing-body combination at the same axial stations. On the basis of the conclusion just stated, the drag rise for this body of revolution should be similar to that for the wing-body combination.

Whitcomb, Richard T.↗
NASA NTRS1981

A review of flight-to-wind tunnel drag correlation

Comparisons are made of wind-tunnel-model and flight drag data for various configurations representing aircraft from the mid-1940s to the 1970s. Discrepancies between model and flight data such as Reynolds number effects, wall interference, and aeroelastic problems are discussed. String support effects and the inability of models to simulate surface deflections for longitudinal trim are also studied. A wind tunnel-to-flight correlation of turbulent friction drag confirms the incompressible Karman-Schoenherr variation of turbulent skin friction with Reynolds number and the T' method for accounting compressibility effects. NASA tested 10 deg cone research indicates that model tests which are affected by tunnel noise may require the lower disturbance level environment available in flight, and it is concluded that new cryogenic facilities will improve the fidelity of model simulations of full-scale flight flow phenomena.

Saltzman, E. J.↗
NASA NTRS1984

Embedded shear layer computations for increased drag reduction

One of the most promising methods of minimizing drag is the reduction of skin friction by injection of low momentum fluid into the near-wall region of turbulent boundary layer flows. This method could be made more effective by limiting the spread rate of the resulting mixing region. In order to achieve a better understanding of how this goal might be achieved, numerical investigations of the relevant fluid dynamic processes governing these regions have been conducted. A compact finite-difference algorithm has been applied to the complete form of the governing conservation equations for a two-dimensional laminar mixing layer. The ability of this computational approach to model successfully the formation and interaction of the large scale vortical structures which dominate such flow fields is verified in the present study. Parameters which affect the spread rate of the mixing region are also identified. In addition, the relative importance of viscous and momentum transport effects in the vortex interactions is determined.

Gatski, T. B.↗
NASA NTRS1988

The fix for tough spots

Modern turbulence-control techniques are discussed. Particular atention is given to retrofit techniques such as riblets and large-eddy breakup (LEBU) devices which use passive elements suitable for a variety of existing vehicles with minimum added complexity. Riblets are small flow-aligned grooves in the aircraft skin that damp turbulence and reduce skin friction; the mechanism of riblet drag reduction derives from the enhancement of turbulence-altering, transverse viscous forces by strong spanwise surface geometry gradients. LEBUs are thin plates or ribbons suspended in a turbulent boundary layer to sever or break up the large vortices that form the convoluted outer edge of the layer. Other turbulence-control techniques are discussed, including one that involves the injection of control vortices into the turbulent boundary layer to modify or substitute for large-eddy structures.

Anders, John B.↗
NASA NTRS1994

A comparison of turbulence models in computing multi-element airfoil flows

Four different turbulence models are used to compute the flow over a three-element airfoil configuration. These models are the one-equation Baldwin-Barth model, the one-equation Spalart-Allmaras model, a two-equation k-omega model, and a new one-equation Durbin-Mansour model. The flow is computed using the INS2D two-dimensional incompressible Navier-Stokes solver. An overset Chimera grid approach is utilized. Grid resolution tests are presented, and manual solution-adaptation of the grid was performed. The performance of each of the models is evaluated for test cases involving different angles-of-attack, Reynolds numbers, and flap riggings. The resulting surface pressure coefficients, skin friction, velocity profiles, and lift, drag, and moment coefficients are compared with experimental data. The models produce very similar results in most cases. Excellent agreement between computational and experimental surface pressures was observed, but only moderately good agreement was seen in the velocity profile data. In general, the difference between the predictions of the different models was less than the difference between the computational and experimental data.

Rogers, Stuart E.↗
NASA NTRS1981

Comparisons of AEROX computer program predictions of lift and induced drag with flight test data

The AEROX aerodynamic computer program which provides accurate predictions of induced drag and trim drag for the full angle of attack range and for Mach numbers from 0.4 to 3.0 is described. This capability is demonstrated comparing flight test data and AEROX predictions for 17 different tactical aircraft. Values of minimum (skin friction, pressure, and zero lift wave) drag coefficients and lift coefficient offset due to camber (when required) were input from the flight test data to produce total lift and drag curves. The comparisons of trimmed lift drag polars show excellent agreement between the AEROX predictions and the in flight measurements.

Axelson, J.↗
NASA NTRS2013

CFL3D, FUN3d, and NSU3D Contributions to the Fifth Drag Prediction Workshop

Results presented at the Fifth Drag Prediction Workshop using CFL3D, FUN3D, and NSU3D are described. These are calculations on the workshop provided grids and drag adapted grids. The NSU3D results have been updated to reflect an improvement to skin friction calculation on skewed grids. FUN3D results generated after the workshop are included for custom participant generated grids and a grid from a previous workshop. Uniform grid refinement at the design condition shows a tight grouping in calculated drag, where the variation in the pressure component of drag is larger than the skin friction component. At this design condition, A fine-grid drag value was predicted with a smaller drag adjoint adapted grid via tetrahedral adaption to a metric and mixed-element subdivision. The buffet study produced larger variation than the design case, which is attributed to large differences in the predicted side-of-body separation extent. Various modeling and discretization approaches had a strong impact on predicted side-of-body separation. This large wing root separation bubble was not observed in wind tunnel tests indicating that more work is necessary in modeling wing root juncture flows to predict experiments.

Park, Michael A.↗