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NASA NTRS1976

Shockless airfoils with thicknesses of 20.6 and 20.7 percent chord analytically designed for a Mach number of 0.68 and a lift coefficient of 0.40

A 20.8 percent-thick airfoil shape was designed to have shockless inviscid flow at a Mach number of 0.68 and a lift coefficient of 0.40. In order to determine the actual airfoils which would yield this same shockless flow when viscous effects are included, boundary layer displacement thicknesses were subtracted from the inviscid shape for Reynolds numbers of 100 and 35 million. This process yielded airfoils with thicknesses of 20.7 and 20.6 percent, respectively. Subtraction of boundary layer displacement thicknesses for Reynolds numbers below 35 million yielded nonphysical airfoils, that is airfoils with negative thicknesses near tHe trailing edge. The pitching moment about the quarter-chord point at the design condition was -0.082 for the inviscid shape and, consequently, for both airfoils. Off-design calculations for the two airfoils were made using a computer program which provides for the interaction of the inviscid flow and boundary layer solutions. The pressure distributions of the airfoils were shockless for conditions from the design point to lower Mach numbers and lift coefficients. No boundary layer separation was predicted except in the last 3 percent chord on the upper surface.

Allison, D. O.↗

NASA NTRS2015

Numerical Calibration of Mass Flow Plug for Inlet Testing

A simple control volume model has been developed to calculate the discharge coefficient through a mass flow plug (MFP) and validated with a calibration experiment. The maximum error of the model within the operating region of the MFP is 0.54%. The control volume analysis developed work is comprised of a sequence of flow calculations through the MFP. The model uses the MFP geometry and operating pressure and temperature to couple continuity, momentum, energy, an equation of state, and wall shear. The discharge coefficient calculation also includes the effects of boundary layer growth, including the reduction in cross-sectional flow area as characterized by the boundary layer displacement thickness. The last calculation in the sequence uses an integral method to calculate the growth of the boundary layer, from which the displacement thickness is then determined. The result of these successive calculations is an accurate one-dimension model of the velocity, pressure, and temperature through the MFP. For comparison, a computational fluid dynamic (CFD) calibration is shown, which when compared to the presented numerical model, had a lower accuracy with a maximum error of 1.35% in addition to being slower by a factor of 100."

Flow Measurement↗

NASA NTRS1986

Application of viscous-inviscid interaction methods to transonic turbulent flows

Two different viscous-inviscid interaction schemes were developed for the analysis of steady, turbulent, transonic, separated flows over axisymmetric bodies. The viscous and inviscid solutions are coupled through the displacement concept using a transpiration velocity approach. In the semi-inverse interaction scheme, the viscous and inviscid equations are solved in an explicitly separate manner and the displacement thickness distribution is iteratively updated by a simple coupling algorithm. In the simultaneous interaction method, local solutions of viscous and inviscid equations are treated simultaneously, and the displacement thickness is treated as an unknown and is obtained as a part of the solution through a global iteration procedure. The inviscid flow region is described by a direct finite-difference solution of a velocity potential equation in conservative form. The potential equation is solved on a numerically generated mesh by an approximate factorization (AF2) scheme in the semi-inverse interaction method and by a successive line overrelaxation (SLOR) scheme in the simultaneous interaction method. The boundary-layer equations are used for the viscous flow region. The continuity and momentum equations are solved inversely in a coupled manner using a fully implicit finite-difference scheme.

Lee, D.↗

NASA NTRS1975

Wind tunnel investigation of a Centaur standard shroud compartment vent from Mach number of 0.70 to 1.96

An experimental investigation was conducted in the Lewis Research Center 8- by 6-foot supersonic wind tunnel to determine the vent discharge coefficient for the Centaur standard shroud/liquid hydrogen tank compartment vent. The test was conducted from Mach 0.70 to 1.96 with the vent mounted in a flat plate. Full scale simulated flight hardware, such as the vent, corrugations, aft field joint ring and ice bag clip was used. Air was discharged from a plenum chamber, located on the tunnel sidewall behind the plate, through five 6.35 cm diameter vent orifices into the free stream. Boundary layer thickeners, analytically predicted displacement thickness for the vehicle nominal flight trajectory could be simulated over the Mach number range. The highest vent discharge coefficient for any given Mach number and vent pressure ratio generally occurred at the maximum displacement thickness.

Johns, A. L.↗

NASA NTRS1983

Self-sustained oscillations of a shock wave interacting with a boundary layer on a supercritical airfoil

A theory is proposed of the self-sustaining oscillations of a weak shock on an airfoil in steady, transonic flow. The interaction of the shock with the boundary layer on the airfoil produces displacement thickness fluctuations which convect downstream and generate sound by interaction with the trailing edge. A feedback loop is established when this sound impinges on the shock wave, resulting in the production of further fluctuations in the displacement thickness. The details are worked out for an idealized mean boundary layer velocity profile, but strong support for the basic hypotheses of the theory is provided by a comparison with recent experiments involving the generation of acoustic "tone bursts' by a supercritical airfoil section.

Ventres, C. S.↗

NASA NTRS1984

Self-sustained oscillations of a shock wave interacting with a boundary layer on a supercritical airfoil

A theory is proposed of the self-sustaining oscillations of a weak shock on an airfoi in steady, transonic flow. The interaction of the shock with the boundary layer on the airfoil produces displacement thickness fluctuations which convect downstream and generate sound by interaction with the trailing edge. A feedback loop is established when this sound impinges on the shock wave, resulting in the production of further fluctuations in the displacement thickness. The details are worked out for an idealized mean boundary layer velocity profile, but strong support for the basic hypotheses of the theory is provided by a comparison with recent experiments involving the generation of acoustic 'tone bursts' by a supercritical airfoil section.

Ventres, C. S.↗

NASA NTRS1987

Comparison of two procedures for predicting rocket engine nozzle performance

Two nozzle performance prediction procedures which are based on the standardized JANNAF methodology are presented and compared for four rocket engine nozzles. The first procedure required operator intercedence to transfer data between the individual performance programs. The second procedure is more automated in that all necessary programs are collected into a single computer code, thereby eliminating the need for data reformatting. Results from both procedures show similar trends but quantitative differences. Agreement was best in the predictions of specific impulse and local skin friction coefficient. Other compared quantities include characteristic velocity, thrust coefficient, thrust decrement, boundary layer displacement thickness, momentum thickness, and heat loss rate to the wall. Effects of wall temperature profile used as an input to the programs was investigated by running three wall temperature profiles. It was found that this change greatly affected the boundary layer displacement thickness and heat loss to the wall. The other quantities, however, were not drastically affected by the wall temperature profile change.

Davidian, Kenneth J.↗

NASA NTRS1987

Comparison of two procedures for predicting rocket engine nozzle performance

Davidian, Kenneth J.↗

NASA NTRS1978

Effect of perturbed flow on the transition from the supersonic laminar boundary layer to the turbulent

Results of experimental studies on the effect of various factors on the transition of a supersonic boundary layer are discussed. It is shown that in supersonic wind tunnels, a significant effect on the transition of the boundary layer on a model is exerted by the scale of acoustic perturbations, which is proportional to the boundary layer displacement thickness of the working section. Experimental data obtained over a wide range of variation of flow parameters in aerodynamically similar test installations with different dimensions of the working section are generalized by means of a correlation parameter based on the displacement thickness.

Kharitonov, A. M.↗

NASA NTRS1945

Wind-Tunnel Tests of a 1/5-Scale Semispan Model of the Republic XF-12 Horizontal Tail Surface

Wind-tunnel tests of a 1/5-scale semispan model of the Republic XF-12 horizontal tail surface equipped with an internally balanced elevator were conducted in the 6- by 6-foot test section of the Langley stability tunnel. The tests included measurements of the aerodynamic characteristics of the horizontal tail with and without a beveled trailing edge and also included measurements of the tab characteristics. The variation of the aerodynamic characteristics with boundary-layer conditions and leakage in the internal-balance chambers, measurements of the boundary-layer displacement thickness near the elevator hinge axis, and pressure distributions at the mean geometric chord were also obtained. The results showed that the hinge-moment characteristics of the elevator were critical to boundary-layer conditions and internal-balance leakage. Increasing the boundary-layer displacement thickness by use of roughness strips reduced the rate of change of elevator hinge moments with tab deflection by about 20 percent. The present horizontal tail appears to be unsatisfactory for longitudinal stability with power on, however, an increase in horizontal-tail lift effectiveness should correct this difficulty. The maneuvering stick force per unit acceleration will be extremely critical to minor variations of the elevator hinge moments if the elevator is linked directly to the stick.

Denaci, H. G.↗

NASA NTRS1983

Studies of sidewall boundary layer in the Langley 0.3 meter transonic cryogenic tunnel with and without suction

Boundary layer measurements on the sidewalls of the Langley 0.3 Meter Transonic Cryogenic Tunnel were made to determine the effectiveness of the passive boundary layer bleed system over a Reynolds number range from 20 to 200 x 10 to the sixth power per meter at Mach numbers from 0.30 to 0.76. The tunnel sidewall boundary layer displacement thickness was about 2 percent of the width of the test section without the boundary layer bleed. Measured velocity profiles correlated well with the defect law of Hama. With the boundary layer bleed equivalent to about 2 percent of the test section mass flow, the boundary layer displacement thickness reduced to about 1 percent of the test section width, which is generally considered acceptable for testing airfoils. It was also noticed that effectiveness of the bleed was nearly independent of the Mach number and Reynolds number over the range of conditions tested. A comparison of the measured suction effectiveness of the bleed with the finite difference and integral methods of boundary layer calculation showed good agreement.

Murthy, A. V.↗

NASA NTRS1990

Leading-edge effects on boundary-layer receptivity

Numerical calculations are presented for the incompressible flow over a parabolic cylinder. The computational domain extends from a region upstream of the body downstream to the region where the Blasius boundary-layer solution holds. A steady mean flow solution is computed and the results for the scaled surface vorticity, surface pressure and displacement thickness are compared to previous studies. The unsteady problem is then formulated as a perturbation solution starting with and evolving from the mean flow. The response to irrotational time harmonic pulsation of the free-stream is examined. Results for the initial development of the velocity profile and displacement thickness are presented. These calculations will be extended to later times to investigate the initiation of instability waves within the boundary-layer.

Gatski, Thomas B.↗

NASA NTRS1982

Recent sidewall boundary-layer investigations with suction in the Langley 0.3-m Transonic Cryogenic Tunnel

An experimental and theoretical study of the Langley 0.3-m Transonic Cryogenic Tunnel (TCT) sidewall boundary-layer with and without suction, has been made. Without suction, the boundary-layer displacement thickness at a station ahead of the model varied from about 1.6 mm to 1.3 mm over a Reynolds number range of 20 million to 200 million per m at Mach numbers from 0.30 to 0.76. Measured velocity profiles were correlated using the defect law of Hama. The boundary-layer displacement thickness decreased when suction was applied; however, after suction of about 2 percent of test section mass flow, the change in the thickness was small. A comparison of the measured suction effectiveness with finite difference and integral methods of boundary-layer calculation showed that both methods predicted the right trend over the range of suction velocities (up to a suction to free-stream velocity ratio of -0.02).

Murthy, A. V.↗

NASA NTRS1984

Applicability of a panel method, which includes nonlinear effects, to a forward-swept-wing aircraft

The ability of a lower order panel method VSAERO, to accurately predict the lift and pitching moment of a complete forward-swept-wing/canard configuration was investigated. The program can simulate nonlinear effects including boundary-layer displacement thickness, wake roll up, and to a limited extent, separated wakes. The predictions were compared with experimental data obtained using a small-scale model in the 7- by 10- Foot Wind Tunnel at NASA Ames Research Center. For the particular configuration under investigation, wake roll up had only a small effect on the force and moment predictions. The effect of the displacement thickness modeling was to reduce the lift curve slope slightly, thus bringing the predicted lift into good agreement with the measured value. Pitching moment predictions were also improved by the boundary-layer simulation. The separation modeling was found to be sensitive to user inputs, but appears to give a reasonable representation of a separated wake. In general, the nonlinear capabilities of the code were found to improve the agreement with experimental data. The usefullness of the code would be enhanced by improving the reliability of the separated wake modeling and by the addition of a leading edge separation model.

Ross, J. C.↗

NASA NTRS1984

Comparison of theoretical and flight-measured local flow aerodynamics for a low-aspect-ratio fin

Flight test and theoretical aerodynamic data were obtained for a flight test fixture mounted on the underside of an F-104G aircraft. The theoretical data were generated using two codes, a two dimensional transonic code called Code H, and a three dimensional subsonic and supersonic code call wing-body. Pressure distributions generated by the codes for the flight test fixture as well as boundary layer displacement thickness generated by the two dimensional code were compared to the flight test data. The two dimensional code pressure distributions compared well except at the minimum pressure point and trailing edge. Shock locations compared well except at high transonic speeds. The three dimensional code pressure distributions compared well except at the trailing edge of the flight test fixture. The two dimensional code does not predict displacement thickness of the flight test fixture well.

Johnson, J. B.↗

NASA NTRS1975

Solutions for incompressible separated boundary layers including viscous-inviscid interaction

Numerical solutions are presented for the laminar and turbulent boundary-layer equations for incompressible flows with separation and reattachment. The separation angularity is avoided by using an inverse technique in which the displacement thickness is prescribed and the pressure is deduced from the resulting solution. The turbulent results appear qualitatively correct despite the use of a two-layer eddy-viscosity model which is generally assumed appropriate only for mild-pressure-gradient flows. A new viscous-inviscid interaction technique is presented in which the inviscid flow is solved inversely by prescribing the pressure from the boundary-layer solution and deducing the new displacement thickness from the solution of a Cauchy integral. Calculations are presented using this interaction procedure for a laminar flow in which separation and reattachment occur on a solid surface.

Carter, J. E.↗

NASA NTRS1978

A viscous-inviscid interactive compressor calculations

A viscous-inviscid interactive procedure for subsonic flow is developed and applied to an axial compressor stage. Calculations are carried out on a two-dimensional blade-to-blade region of constant radius assumed to occupy a mid-span location. Hub and tip effects are neglected. The Euler equations are solved by MacCormack's method, a viscous marching procedure is used in the boundary layers and wake, and an iterative interaction scheme is constructed that matches them in a way that incorporates information related to momentum and enthalpy thicknesses as well as the displacement thickness. The calculations are quasi-three-dimensional in the sense that the boundary layer and wake solutions allow for the presence of spanwise (radial) velocities.

Johnston, W.↗

NASA NTRS1978

A viscous-inviscid interactive compressor calculation

A viscous-inviscid interactive procedure for subsonic flow is developed and applied to an axial compressor stage. Calculations are carried out on a two-dimensional blade-to-blade region of constant radius assumed to occupy a mid-span location. Hub and tip effects are neglected. The Euler Equations are solved by MacCormack's method, a viscous marching procedure is used in the boundary layers and wake, and an iterative interaction scheme is constructed that matches them in a way that incorporates information related to momentum and enthalpy thicknesses as well as the displacement thickness. The calculations are quasi-three-dimensional in the sense that the boundary layer and wake solutions allow for the presence of spanwise (radial) velocities.

Johnston, W. A.↗