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The period of the large coherent structure in a subsonic, compressible, turbulent boundary layer was determined using the autocorrelation of the velocity and pressure fluctuations for Reynolds numbers between 5,000 and 35,000. In low Reynolds number flows the overall correlation period scaled with the outer variables - namely, the free stream velocity and the boundary layer thickness.
The static aerodynamic characteristics were studied on a model wing-body concept for a high-speed research airplane in a low-turbulence pressure tunnel. The experiment consisted of configuration buildup from the basic body by adding a wing, center vertical tail, three-module scramjet, and six-module scramjet engine. The test Mach number was 0.2 at Reynolds numbers, based on fuselage length, ranging from 2.78 x 1 million to 23 x 2 million. The test angle-of-attack range was approximately -5 to 30 deg at constant angles of sideslip of 0 deg and 4 deg. The elevons were deflected from 5 deg to -15 deg. Roll and yaw control were investigated.
A Au aircraft model with a close-coupled canard mounted above the wing chord plane was considered. Model angle of attack was varied from -4 deg to 15 deg; canard incidence was varied from -5 deg to 18 deg; and selected canard and wing flap deflections were investigated. By using the canard incidence for trim, maximum trimmed lift-drag ratios of about 8.8, 7.7, and 4.7 were obtained at free-stream Mach numbers of 0.40, 0.90, and 1.20, respectively. At a lift coefficient of 0.60, model trim angle of attack could be varied over an incremental range between 3.0 deg and 3.8 deg, depending on Mach number, by different combinations of control settings. At high lift coefficients, larger trimmed lift-drag ratios were obtained by using the deflection capability of the canard leading- and trailing-edge flaps before increasing canard incidence angle.
The induced lift characteristics are studied for a vectored-thrust concept in which jet-exhaust nozzles were located in the fuselage at the wing trailing edge. The wing had a supercritical airfoil section. The investigation was conducted at Mach numbers from 0.4 to 0.95, angles of attack up to 14 deg, and thrust coefficients up to 0.35, and nozzle deflection angles of 0 deg and 30 deg. Separate force balances were used to determine both total aerodynamic and thrust forces alone which allowed for a direct measurement of jet turning angle at forward speeds. The Reynolds number per meter varied from 8.20 x 1 million to 12.80 x 1 million. The results show that the configuration with the supercritical wing generally had a better performance with respect to both lift augmentation and drag reduction than the same configuration with a 64 series airfoil.
The characteristics of a thick hypersonic boundary layer turbulent for a length of 175 cm on a 4 deg sharp wedge were measured. The resulting boundary layer was free from transverse curvature effects and only mildly affected by upstream history effects caused by pressure and wall temperature gradients. Heat-transfer distributions were used to locate regions of laminar, transitional, and turbulent flow at an edge unit Reynolds number of 470,000 cm at wall-to-total temperature ratios from about 0.3 to 1. Wall cooling had little effect on the location of the transition region. Pitot and total temperature profiles and skin-friction measurements were obtained at several locations along the model longitudinal centerline. Mixing length and turbulent Prandtl number distributions were derived from the fully turbulent mean profiles.
To extrapolate from subscale wind tunnel tests to full-scale flight is a well-recognized problem. It becomes especially critical for present day high performance missiles and aircraft which operate at high angles of attack where separated flow effects often dominate the vehicle aerodynamics. The dynamic effects of flow separation can usually not be simulated at subscale Reynolds numbers. This is illustrated by examples of tests with slender vehicles and high performance aircraft. It is shown how analytic means can be developed establishing theoretical relationships between dynamic and static aerodynamic characteristics and how such means make it possible to extrapolate analytically from subscale tests to full-scale flight. The roll of future high Reynolds number facilities in establishing such 'analytic extrapolation' tools is discussed.
An equation in the form of a Fourier series was derived in an earlier study to give displacements of the originally straight edge of a plate as the result of a small heat input patch moving at speed c along that edge. That unwieldly formulation is reduced to a simple integral equation, which is shown to be manageable in an example calculation. The integral equation is shown to be closely related to that of Ling and Mow, although it is much simpler to use. Comparison of the results of the simplification to results of the original series formulation suggest that the equation proposed for 'high Peclet number' may be a valid approximation for Peclet number as low as 2.5.
Computations, employing several turbulence models, are compared with a series of attached supersonic turbulent boundary-layer experiments over an extensive range of Reynolds numbers (11.7 x 10 to the 6th to 314 x 10 to the 6th). These experiments included measurements of surface pressure and skin friction for adverse pressure gradients ranging in magnitude from those of previous investigations to an order of magnitude greater. The turbulence models evaluated include algebraic and two-equation eddy-viscosity models and two full Reynolds stress models. In general all the models tested performed well independently of the magnitude of the pressure gradient or Reynolds number and could predict the measured skin friction for most cases with sufficient accuracy for engineering purposes.
An experimental study has been carried out to detail the interaction of a compressible turbulent boundary layer with shock waves of varying strengths. The interaction was produced by two-dimensional compression corners of 8, 16, 20, and 24 deg angles. The incoming boundary layer had an edge Mach number of 2.85 and a Reynolds number of 1.7 million based on overall thickness. Detailed mean-flow and surface measurements are presented for the four corner angles. The 8 deg corner flow was found to be fully attached, while the 16 deg case was near incipient separation. Both the 20 deg and 24 deg corners produced significant flow separation regions. In the discussion of these results, emphasis is placed on the development of flowfield properties from attached to separated conditions. Comparisons made with a computational solution of the Navier-Stokes equations show good agreement when the corner flow is not separated.
The problem of inferring the speed of sound in a contained spherically symmetric fluid solely from its natural frequencies of vibration is considered. An investigation of the case in which the data consist of the two spectra associated with the angular numbers 0 and 1, suggests the possibility that a one-parameter family of slowness profiles can be constructed. These profiles are compatible with the data, up to first order in the non-uniformity of the fluid. It is conjectured that for other angular numbers, the loss of information increases as the difference between them increases.
A review is presented of implicit and hybrid methods applicable to solving viscous flows at high Reynolds numbers. Flows within axisymmetric channels containing stationary shock waves, past blunt-nosed lifting airfoils, past sharp-nosed symmetric airfoils with buffet, past three-dimensional compression ramps with side walls, and past ogive- and hemisphere-cylinders at angle of attack have been examined. Reynolds numbers as high as 10 to the 9th power have been used.
Performance analysis of a gas-driven shock tube shows that transonic airfoil flows with chord Reynolds numbers of the order of 100 million can be produced, with limitations being imposed by the structural integrity of the facility or the model. A study of flow development over a simple circular arc airfoil at zero angle of attack was carried out in a shock tube at low and intermediate Reynolds numbers to assess the testing technique. Results obtained from schlieren photography and airfoil pressure measurements show that steady transonic flows similar to those produced for the same airfoil in a wind tunnel can be generated within the available testing time in a shock tube with properly contoured test section walls.
The results of the study reveal that with regard to the pulmonary lesions, twice the number of exposures is compensated for by quartering the overpressure of the wave crest. With regard to the mortality rates, it reveals that halving the overpressure of the wave crest is offset by a 20-fold increase in the number of exposures.
An experimental aerodynamic heating investigation was conducted to determine effects of hot boundary-layer ingestion into the cove on the windward surface between a wing and elevon for cove seal leak areas nominally between 0 and 100 percent of cove entrance area. Pressure and heating-rate distributions were obtained on the wing and elevon surfaces and on the cove walls of a full-scale model that represented a section of the cove region on the space shuttle orbiter. Data were obtained for both attached and separated turbulent boundary layers upstream of the unswept cove entrance. Average free-stream Mach number was 6.9, average free-stream unit Reynolds numbers were 1.31 x 10 to the 6th power and 4.40 x 10 to the 6th power per meter (0.40 x 10 to the 6th power and 1.34 x 10 to the 6th power per foot), and average total temperature was 1888 K (3400 R). Cove pressures and heating rates varied as a function of seal leak area independent of leak aspect ratio. Although cove heating rates for attached flow did not appear intolerable, it was postulated that convective heating in the cove may increase with time. For separated flow, the cove environment was considered too severe for unprotected interior structures of control surfaces.
Tabulated surface pressure data for a series of forebodies which have analytically defined cross sections and are based on a 20 degs half-angle cone are presented without analysis. Five of the cross sections were ellipses having axis ratios of 3/1, 2/1, 1/1, 1/2, and 1/3. The sixth cross section was defined by a curve having a single lobe. The data generally cover angles of attack from -5 degs to 20 degs at angles of sideslip from 0 degs to 5 degs for Mach numbers of 1.70, 2.50, 3.95, and 4.50 at a constant Reynolds number.
The tabulated results of surface pressure tests conducted on the wing and fuselage of an airplane model in the Langley Unitary Plan wind tunnel are presented without analysis. The model tested was that of a supersonic-cruise airplane with a highly swept arrow-wing planform, two engine nacelles mounted beneath the wing, and outboard vertical tails. Data were obtained at Mach numbers of 2.30, 2.96, and 3.30 for angles of attack from -4 deg to 12 deg. The Reynolds number for these tests was 6,560,000 per meter.
A short, annular dump diffuser with suction stabilized vortices in the region of abrupt area change was tested with a full scale, annular swirl can combustor. The prediffuser area ratio was 1.4. Performance data were obtained for both isothermal and burning conditions at inlet temperatures of 589 to 895 K and pressures of 0.5 to 1.0 MPa for a range of diffuser inlet Mach numbers from 0.25 to 0.53. Suction rates were 0 to 20 percent of the total diffuser mass flow rate. Diffuser effectiveness increased from 47 percent without suction to approximately 80 percent for a total suction rate of 14 percent. Combustor total pressure loss for the same total suction rate was reduced from 6.8 percent without suction to 4.0 percent at an inlet Mach number of 0.40.
A large amount of fundamental aerodynamic missile data, which were stored for a number of years at the Applied Physics Laboratory, are reported. Data that supplements the M = 2.0 flow field data are provided. The Mach number effect by means of pressure fields only, at M = 1.5 and 2.0, and at angles of attack up to 23 deg at a mid-body station where a wing might be located is described.