Showing papers on "Drag divergence Mach number published in 1987"

Experimental investigation of heterogeneous compressible shear layers

Abstract: The compressible, two-dimensional shear layer is investigated experimentally in a novel facility. In this facility, it is possible to flow similar or, dissimilar gases of different densities and to select different Mach numbers for each stream over a wide range of Reynolds numbers. In the current experiments, ten combinations of gases and Mach numbers are studied in which the freestream Mach numbers range from 0.2 to 4, the density ratio varies from 0.2 to 9.2, and the velocity ratio varies from 0.13 to 1. The growth of the turbulent region of the layer is measured by means of pitot pressure profiles obtained at several streamwise locations. The resulting growth rate is estimated to be about 80% of the visual growth rate. The transition from laminar to turbulent flow, as well as the structure of the turbulent layer, are observed with Schlieren photographs of 20 nanosecond duration. Streamwise pressure distribution and total pressures are measured by means of a Scanivalve-pressure transducer system. An underlying objective of this investigation was the definition of a compressibility-effect parameter that correlates and consolidates the experimental results, especially the turbulent growth rates. A brief analytical investigation of the vortex sheet suggests that such a parameter is the Mach number in a frame of reference moving with the phase speed of the disturbance, called here the convective Mach number. In a similar manner, the convective Mach number of a turbulent shear layer is defined as the one seen by an observer moving with the convective velocity of the dominant waves and structures. It happens to have about the same value for each stream. In the current experiments, it ranges from 0 to 1.9. The correlations of the growth rate with convective Mach number fall approximately onto one curve when the growth rate is normalized by its incompressible value at the same velocity and density ratios. The normalized growth rate, which is unity for incompressible flow, decreases gradually with increasing convective Mach number, reaching an asymptotic value of about 0.25 for supersonic convective Mach numbers. The above behavior is in qualitative agreement with results of linear stability theory as well as with those of previous, one-stream experiments. Large-scale structures, resembling those observed in subsonic shear layers, are evident in the Schlieren photographs. It is estimated that the mean structure spacing, normalized by the local thickness, is reduced to about half its incompressible value as the convective Mach number becomes supersonic. An estimate of the transition Reynolds number has been obtained from the photographs of two shear layers having quite different convective Mach numbers, one low subsonic and the other sonic. In both cases, it is about 2 x 105, based on distance to transition and properties of the high unit Reynolds number stream, thus suggesting that, in this experiment, transition is dominated by instabilities of the wake, rather than of the shear layer.

Viscous transonic airfoil workshop results using ARC2D

Abstract: Computations have been performed in response to the Viscous Transonic Airfoil Workshop associated with the AIAA 25th Aerospace Sciences Meeting (January 1987). The purpose of the workshop is to establish the capabilities of various methods for computing viscous flowfields for a range of conditions and geometries. The results of the test cases will demonstrate the capabilities of the methods in predicting both aerodynamic trends and flowfield details. ARC2D, a well-established Navier-Stokes code, was used to compute the flowfields for the designated airfoils, Mach numbers, angles of attack and other specifications of the Workshop committee.

Design of the low-speed NLF(1)-0414F and the high-speed HSNLF(1)-0213 airfoils with high-lift systems

Abstract: The design and testing of Natural Laminar Flow (NLF) airfoils is examined. The NLF airfoil was designed for low speed, having a low profile drag at high chord Reynolds numbers. The success of the low speed NLF airfoil sparked interest in a high speed NLF airfoil applied to a single engine business jet with an unswept wing. Work was also conducted on the two dimensional flap design. The airfoil was decambered by removing the aft loading, however, high design Mach numbers are possible by increasing the aft loading and reducing the camber overall on the airfoil. This approach would also allow for flatter acceleration regions which are more stabilizing for cross flow disturbances. Sweep could then be used to increase the design Mach number to a higher value also. There would be some degradation of high lift by decambering the airfoil overall, and this aspect would have to be considered in a final design.

Multi-body aircraft with an all-movable center fuselage actively controlling fuselage pressure drag

Abstract: A multi-body aircraft with an all-movable center fuselage 20 which translates relative to two side fuselages 22. At subsonic and transonic flight the center fuselage 20 is in a forward position. At supersonic speeds the center fuselage 20 moves aft so as to ensure optimum aerodynamic interference at particular Mach numbers. This provides an increased shock strength and greater surface area so that significant reductions in zero-lift wave drag can be achieved. This concept allows for a significant increase in the wing aspect ratio which would improve high-lift performance at all speed without incurring a significant supersonic zero-lift wave drag penalty. In addition, an improved low-fineness ratio, high-speed performance is achieved at all speeds and for all flight conditions.

High Reynolds number transonic tests on a NACA 0012 airfoil in the Langley 0.3-meter transonic cryogenic tunnel

Abstract: Tests were conducted in the two-dimensional test section of the Langley 0.3-m Transonic Cryogenic Tunnel on a NACA 0012 airfoil to obtain aerodynamic data as a part of the Advanced Technology Airfoil Test (ATAT) program. The test program covered a Mach number range of 0.30 to 0.82 and a Reynolds number range of 3.0 to 45.0 x 10 to the 6th. The stagnation pressure was varied between 1.2 and 6.0 atmospheres and the stagnation temperature was varied between 300 K and 90 K to obtain these test conditions. Plots of the spanwise variation of drag coefficient as a function of normal force coefficient and the variation of the basic aerodynamic characteristics with angle of attack are shown. The data are presented uncorrected for wall interference effects and without analysis.

Dynamics of Particle—Shock Interactions: Part I: Similitude

Abstract: The equation of particle motion is normalized for either relaxation behind a shock front or particle impaction from supersonic streams. Algebraic expressions are introduced for the drag coefficient in both cases to account for non-Stokesian drag and free molecular flow. Expressions are developed for the particle Knudsen, Reynolds, Mach, and Stokes numbers and for the ratio of the local gas-to-particle density and maximum particle drag. It is demonstrated that knowledge of the particle stagnation Knudsen number, local gas Mach number, and specific heat ratio are sufficient to predict the particle Knudsen, Reynolds, and Stokes numbers and the maximum drag behind the shock front. Graphs are also provided to estimate these important particle parameters.

An experimental investigation of a supercritical airfoil at transonic speeds

Abstract: Benchmark experimental data obtained in the two-dimensional, transonic flow field surrounding a supercritical airfoil are presented. Airfoil surface and tunnel wall pressure and LDV measurements are used to describe the flow on the model, above the wing and in the wake. Comparisons are made with calculations using the Reynolds-averaged Navier-Stokes equations. The results illustrate the performance of two turbulence models in both separated and attached flows. The largest differences between theory and experiment occurred in separated flows with the Johnson and King turbulence model providing the best estimates.

Aerodynamic Effects of Probe-Induced Flow Separation on Bluff Bodies at Transonic Mach Numbers

Abstract: An experimental investigation using probe-induced flow separation devices to reduce transonic blunt-body drag has been conducted. Particularly examined were blunt axisymmetric mainbodies with axially aligned cylindrical probes extended ahead. The experiments were performed in a ballistic facility, and the data obtained include drag coefficients and shadowgraphs of the flowfield. Drag reductions of 25% were observed over the Mach number range 0.85-1.25. Flow visualization reveals a distinct difference in the manner in which the flow reattaches to the mainbody for lowand high-drag geometries. The flowfield also exhibited modes resembling open- and closed-cavity flows, depending primarily on the probe length. Large-scale flow oscillations were observed for both low- and high-drag cases. This work clearly demonstrates the possibility for transonic drag reductions using cylindrical probes and provides useful information on more fundamental questions concerning transonic flow separation and reattachment.

Low-speed wind-tunnel results for symmetrical NASA LS(1)-0013 airfoil

Abstract: A wind-tunnel test has been conducted in the Langley Low-Turbulence Pressure Tunnel to evaluate the performance of a symmetrical NASA LS(1)-0013 airfoil which is a 13-percent-thick, low-speed airfoil. The airfoil contour was obtained from the thickness distribution of a 13-percent-thick, high-performance airfoil developed for general aviation airplanes. The tests were conducted at Mach numbers from 0.10 tp 0.37 over a Reynolds number range from about 0.6 to 12.0 X 10 to the 6th power. The angle of attack varied from about -8 to 20 degrees. The results indicate that the aerodynamic characteristics of the present airfoil are similar to, but slightly better than, those of the NACA 0012 airfoil.

An Examination of Controlled Vortex Drag Using Stepped Afterbodies from M = 0.5 to 3.0

Abstract: : A comparison of the zero lift drag coefficients of a stepped base projectile to flat base and truncated boattail base projectiles is presented. Three model configurations were investigated during the test program. These included an experimental 20mm round with a 7 1/2 deg., truncated boattail base, a round modified with a flat base, and a round modified with a stepped base. All of the projectiles were tested at sea level conditions in an indoor ballistic free-flight facility. This paper discusses the aerodynamic experiment and the data obtained. Results show that the zero lift drag coefficient of the stepped base projectile was less than that of the flat base round for the subsonic Mach number range and approximately the same for the transonic and supersonic ranges. However, the stepped base projectile produced zero lift drag greater than that of the boattail round at each Mach number. Keywords: Base drag; Separated flow; projectile afterbody; Ballistic testing.

Afterbody flowfield computations at transonic and supersonic Mach numbers

Abstract: The flowfield around various axisymmetric afterbody configurations is computed with a finite-volume NavierStokes code, incorporating a total variation diminishing implicit upwind-biased scheme for high accuracy and using alternatively the k-e or the Baldwin-Lomax turbulence model Computations are done for both solid plume simulators and real jet flows Results for two geometries at several combinations of jet and freestream conditions are shown Agreement with the experimental data is very good

Pressure field due to drag reducing outer layer devices in turbulent boundary layers

Abstract: The wall static pressure in the vicinity of drag reducing outer layer devices in flat wall turbulent boundary layers has been measured and compared with an inviscid theory. Symmetric and cambered airfoil devices have been examined at small angles of attack and very low chord Reynolds numbers. Airfoil devices impose a sequence of strong favorable and adverse pressure gradients on the boundary layer whose drag is to be reduced. At very small angles of attack (± 2°), this pressure field extends up to about three chord lengths downstream of the trailing edge of an airfoil device. Also examined are the pressures on the upper and lower surfaces of a symmetric airfoil device in the freestream and near the wall. The freestream pressure distribution around an airfoil section is altered by the wall proximity. The relevance of lift enhancement caused by wall proximity to drag reduction has been discussed. The pressure distributions on the flat wall beneath the symmetric airfoil devices are predicted well by the inviscid theory. However, the remaining pressure distributions are predicted only qualitatively, presumably because of strong viscous effects.

Effects of afterbody boattail design and empennage arrangement on aeropropulsive characteristics of a twin-engine fighter model at transonic speeds

Abstract: The effects of empennage arrangement and afterbody boattail design of nonaxisymmetric nozzles on the aeropropulsive characteristics of a twin-engine fighter-type model have been determined in an investigation conducted in the Langley 16-Foot Transonic Tunnel. Three nonaxisymmetric and one twin axisymmetric convergent-divergent nozzle configurations were tested with three different tail arrangements: a two-tail V-shaped arrangement; a staggered, conventional three-tail arrangement; and a four-tail arrangement similar to that on the F-18. Two of the nonaxisymmetric nozzles were also vectorable. Tests were conducted at Mach numbers from 0.60 to 1.20 over an angle-of-attack range from -3 deg to 9 deg. Nozzle pressure ratio was varied from 1 (jet off) to approximately 12, depending on Mach number. Results indicate that at design nozzle pressure ratio, the medium aspect ratio nozzle (with equal boattail angles on the nozzle sidewalls and upper and lower flaps) had the lowest zero angle of attack drag of the nonaxisymmetric nozzles for all tail configurations at subsonic Mach numbers. The drag levels of the twin axisymmetric nozzles were competitive with those of the medium-aspect-ratio nozzle at subsonic Mach number.

Drag reduction in two-phase gas-liquid flows

Abstract: A considerable amount of drag reduction is observed in various flow regimes for two-phase gas-liquid flow of drag reducing fluids. Various attempts have been made to correlate the amount of drag reduction, but comparison with single phase flow of drag reducing fluids has not been done in a satisfactory manner. It is well known that single phase flow of drag reducing fluids exhibits a diameter effect. In order to account for this, a plot of percent drag reduction versus friction velocity, ..mu../sup */, in the absence of a polymer additive gives a simple method of estimating the amount of drag reduction for a given flow rate and pipe diameter. For two-phase flow no such comparison has been made.

A study of the effects of Reynolds number and Mach number on constant pressure coefficient jump for shock-induced trailing-edge separation

Abstract: The influence of Mach and Reynolds numbers as well as airfoil and planform geometry on the phenomenon of constant shock jump pressure coefficient for conditions of shock induced trailing edge separation (SITES) was studied It was demonstrated that the phenomenon does exist for a wide variety of two and three dimensional flow cases and that the influence of free stream Mach number was not significant The influence of Reynolds number was found to be important but was not strong Airfoil and planform geometric characteristics were found to be very important where the pressure coefficient jump was shown to vary with the sum of: (1) airfoil curvature at the upper surface crest, and (2) camber surface slope at the trailing edge It was also determined that the onset of SITES could be defined as a function of airfoil geometric parameters and Mach number normal to the leading edge This onset prediction was shown to predict the angle of onset to within + or - 1 deg accuracy or better for about 90% of the cases studied

Supersonic Aerodynamic Characteristics of a Space Shuttle Orbiter Model at Angles of Attack from 20 Deg to 90 Deg

Abstract: Space Shuttle Orbiter aerodynamic characteristics have been determined in the angle-of-attack range from 20 to 90 degrees and Mach numbers from 4.60 to 1.80 at Reynolds numbers, based on body length, of 2.15 x 10 to the 6th and 4.30 x 10 to the 6th power. Emphasis is on vehicle stability, control, and trim characteristics above a 60 degree angle of attack. The model used was a 0.986 percent scale Orbiter, having a blade-mounted support system entering the model in the region of the vertical tail. Elevon deflections of 0, -10, -20, and -40 degrees and body-flap deflections of 0, +6 and -12 degrees were investigated individually and in combination. Schlieren photographs are also presented for selected configurations and Mach numbers. The Orbiter was found to be longitudinally stable and trimmable in the angle-of-attack range from approximately 60 to 80 degrees. Both the elevon and body flap provided positive pitch control-effectiveness at angles of attack from 60 to 80 degrees and the Mach numbers of this study. For the range of neutral to stable trim in the angle-of-attack range above about 55 degrees, the deflected elevon/body-flap combination provided positive trimmed lift and lift/drag ratios.

Current CFD efforts in projectile aerodynamics

Abstract: Information is given in viewgraph form on current computational fluid dynamics (CFD) efforts in projectile aerodynamics. Topics covered include spinning projectiles, fin stabilized projectiles, model geometry, the variation of base drag with base bleed, the variation of normal force with Mach number, and chordwise pressure distribution.

Use of LEBU-devices for drag reduction at flight conditions

Abstract: The practical aspects of drag reduction using LEBU (large-eddy break-up) devices, turbulence manipulators, etc. in flight are discussed with the help of experience from previous flight tests. These tests have shown that appreciable reductions in local skin friction exist under flight conditions, and that the turbulence-manipulating effects can be found for swept tandem devices in transonic as well as supersonic flows. The unsteady loads experienced with the devices are discussed in connection with the frequencies occurring in this type of real flight conditions.

Transonic analysis and design of axisymmetric bodies in nonuniform flow

Abstract: An inviscid nonuniform axisymmetric transonic code was developed for applications in analysis and design. Propfan slipstream effect on pressure distribution for a body with and without sting was investigated. Results show that nonuniformity causes pressure coefficient to be more negative and shock strength to be stronger and more rearward. Sting attached to a body reduced the pressure peak and moves the rear shock forward. Extent and Mach profile shapes of the nonuniformity region appeared to have little effect on the pressure distribution. Increasing nonuniformity magnitude made pressure coefficient more negative and moved the shock rearward. Design study was conducted with the CONMIN optimizer for an ellipsoid and a body with the NACA-0012 counter. For the ellipsoid, the general trend showed that to reduce the pressure drag, the front portion of the body should be thinner and the contour of the rear portion should be flatter than the ellipsoid. For the design of a body with a sharp trailing edge in transonic flow with an initial shape given by the NACA-0012 contour, the pressure drag was reduced by decreasing the nose radius and increasing the thickness in the aft portion. Drag reduction percentages are given.

Summary of studies to reduce wing-mounted propfan installation drag on an M = 0.8 transport

Abstract: Powerplant installation losses for an advanced, high-speed, turboprop transport have been investigated in the Ames Research Center Transonic Wind Tunnels as a part of the NASA Advanced Turboprop Program (ATP) Force and pressure tests have been completed at Mach numbers from 06 to 082 on baseline and modified powered-model configurations to determine the magnitude of the losses and to what extent current design tools could be used to optimize the installed performance of turboprop propulsion systems designed to cruise at M = 08 Results of the tests indicate a large reduction in installed drag for the modified configuration The wing-mounted power plant caused destabilizing pitching moments and a negative shift in the zero-lift pitching moment

An aerodynamic model for a hemispherically-capped biconic reentry vehicle with six drag flaps

Abstract: The development of an aerodynamic model for a hemispherically-capped biconic reentry vehicle with six drag flaps is presented. The aerodynamic model is primarily based on wind tunnel test results, with the use of computational fluid dynamic codes. For Mach numbers from 4 to 15, the inviscid axial force coefficient was computed for drag flap deflections from 6 to 36. Axial force coefficient was found to vary significantly with ablating flap shape as well as with changing flight conditions. The aerodynamic model can be used for input to vehicle recovery trajectory simulations.