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NASA NTRS1987

Displacement surface calculations for a hypersonic aircraft

A method is presented to calculate the three-dimensional displacement surface about a hypersonic aircraft. This calculation requires a flow-field-solution to at least the thin-layer Navier-Stokes equations about the aircraft as input. An upwind, finite-volume code developed at NASA Langley Research Center was used to obtain the flow-field solution. The displacement surface is the three-dimensional counterpart to the two-dimensional displacement thickness. Flow-field solutions, along with the displacement surface calculations, are presented for a generic hypersonic aircraft at a Mach number of 24.5 and angle of attack of zero and one degree.

Richardson, Pamela F.↗

NASA NTRS1988

Sidewall boundary-layer measurements with upstream suction in the Langley 0.3-meter transonic cryogenic tunnel

The Langley 0.3 Meter Transonic Cryogenic Tunnel has provision for boundary removal from the sidewalls to reduce sidewall interference effects on the test data. The tests carried out to determine the change in the empty test section sidewall boundary layer thickness at the model station with upstream boundary layer mass removal are described. The boundary layer measurements showed that the upstream removal region is effective in reducing the boundary layer thickness at the model station. The boundary layer displacement thickness reduced from about 1.2 percent to about .4 percent of the test section width. The boundary layer velocity profiles followed a power law variation in the outer region and showed good correlation when plotted in terms of boundary layer momentum thickness.

Murthy, A. V.↗

NASA NTRS1990

A computational efficient modelling of laminar separation bubbles

In predicting the aerodynamic characteristics of airfoils operating at low Reynolds numbers, it is often important to account for the effects of laminar (transitional) separation bubbles. Previous approaches to the modelling of this viscous phenomenon range from fast but sometimes unreliable empirical correlations for the length of the bubble and the associated increase in momentum thickness, to more accurate but significantly slower displacement-thickness iteration methods employing inverse boundary-layer formulations in the separated regions. Since the penalty in computational time associated with the more general methods is unacceptable for airfoil design applications, use of an accurate yet computationally efficient model is highly desirable. To this end, a semi-empirical bubble model was developed and incorporated into the Eppler and Somers airfoil design and analysis program. The generality and the efficiency was achieved by successfully approximating the local viscous/inviscid interaction, the transition location, and the turbulent reattachment process within the framework of an integral boundary-layer method. Comparisons of the predicted aerodynamic characteristics with experimental measurements for several airfoils show excellent and consistent agreement for Reynolds numbers from 2,000,000 down to 100,000.

Dini, Paolo↗

NASA NTRS1980

Theoretical study of nonadiabatic boundary-layer stabilization times in a cryogenic wind tunnel for typical stainless steel wing and fuselage models

The time varying effect of nonadiabatic wall conditions on boundary layer properties was studied for a two dimensional wing section and an axisymmetric fuselage. The wing and fuselage sections are representative of the wing root chord and fuselage of a typical transport model for the National Transonic Facility. The analysis was made with a solid wing and three fuselage configurations (one solid and two hollow with varying skin thicknesses) all made from AISI type 310S stainless steel. The displacement thickness and local skin friction were investigated at a station on the model in terms of the time required for these two boundary layer properties to reach an adiabatic wall condition after a 50 K step change in total temperature. The analysis was made for a free stream Mach number of 0.85, a total temperature of 117 K, and stagnation pressures of 2, 6, and 9 atm.

Johnson, C. B.↗

NASA NTRS1950

The Effect of the Inlet Mach Number and Inlet-boundary-layer Thickness on the Performance of a 23 Degree Conical-diffuser-tail-pipe Combination

An investigation was conducted to determine the effect of the inlet Mach number and entrance-boundary-layer thickness on the performance of a 23 degree 21-inch conical-diffuser - tail-pipe combination with a 2:1 area ratio. The air flows used in this investigation covered an inlet Mach number range from 0.17 to 0.89 and corresponding Reynolds numbers of 1,700,000 to 7,070,000. Results are reported for two inlet-boundary-layer thicknesses. Over the entire range of flows, the mean value of the inlet displacement thickness is about 0.034 inch for the thinner inlet boundary layer and about 0.170 inch for the case of the thicker inlet boundary layer. The performance of the diffuser - tail-pipe combination is presented together with examples of longitudinal static-pressure distribution and the results of boundary-layer pressure surveys made at six points along the diffuser wall. The results indicated a progressive diminution of the static-pressure recovery and a steady increase in the total-pressure losses as the inlet Mach number was increased for both inlet-boundary-layer thicknesses. The ratio of actual static-pressure rise to that theoretically possible was much less and the total-pressure losses were greater for the case of the thicker inlet boundary layer throughout the speed range investigated. With the thinner inlet boundary layer, flow separation occurred at the diffuser exit at all inlet Mach numbers.Unseparated flow alternating with separated flow was observed near the inlet at the higher velocities. For the case of the thicker inlet boundary layer, the origin of the separated region occurred in the vicinity of the inlet-duct-diffuser junction section at all Mach numbers.

Persh, Jerome↗

NASA NTRS2007

Nonlinear Theory and Breakdown

The main points of recent theoretical and computational studies on boundary-layer transition and turbulence are to be highlighted. The work is based on high Reynolds numbers and attention is drawn to nonlinear interactions, breakdowns and scales. The research focuses in particular on truly nonlinear theories, i.e. those for which the mean-flow profile is completely altered from its original state. There appear to be three such theories dealing with unsteady nonlinear pressure-displacement interactions (I), with vortex/wave interactions (II), and with Euler-scale flows (III). Specific recent findings noted for these three, and in quantitative agreement with experiments, are the following. Nonlinear finite-time break-ups occur in I, leading to sublayer eruption and vortex formation; here the theory agrees with experiments (Nishioka) regarding the first spike. II gives rise to finite-distance blowup of displacement thickness, then interaction and break-up as above; this theory agrees with experiments (Klebanoff, Nishioka) on the formation of three-dimensional streets. III leads to the prediction of turbulent boundary-layer micro-scale, displacement-and stress-sublayer-thicknesses.

Smith, Frank↗

NASA NTRS1950

Investigation of High-Subsonic Performance Characteristics of a 12 Degree 21-Inch Conical Diffuser, Including the Effects of Change in Inlet-Boundary-Layer Thickness

Investigations were conducted of a 12 degree 21-inch conical diffuser of 2:l area ratio to determine the interrelation of boundary layer growth and performance characteristics. surveys were made of inlet and exit from, longitudinal static pressures were recorded, and velocity profiles were obtained through an inlet Reynolds number range, determined From mass flows and based on inlet diameter of 1.45 x 10(exp 6) to 7.45 x 10(exp 6) and a Mach number range of 0.11 to approximately choking. These investigations were made to two thicknesses of inlet boundary layer. The mean value, over the entire range of inlet velocities, of the displacement thickness of the thinner inlet boundary layer was approximately 0.035 inch and that of the thicker inlet boundary layer was approximately six times this value. The loss coefficient in the case of the thinner inlet boundary layer had a value between 2 to 3 percent of the inlet impact pressure over most of the air-flow range. The loss coefficient with the thicker inlet boundary layer was of the order of twice that of the thinner inlet boundary layer at low speeds and approximately three times at high speeds. In both cases the values were substantially less than those given in the literature for fully developed pipe flow. The static-pressure rise for the thinner inlet boundary layer was of the order of 95 percent of that theoretically possible over the entire speed range. For the thicker inlet boundary layer the static pressure rise, as a percentage of that theoretically possible, ranged from 82 percent at low speeds to 68 percent at high speeds.

Copp, Martin R.↗

NASA NTRS1981

A viscous-inviscid interaction model of jet entrainment

A viscous-inviscid interaction model for predicting jet entrainment effects on axisymmetric, nozzle afterbodies at subsonic speeds is presented. The model is based on a displacement thickness correction to the inviscid jet boundary that accounts for mixing-induced streamline deflections in the inviscid region. The displacement correction is shown to be related to the local mass entrainment rate and, for thin mixing layers, the model is shown to be analogous to displacement models used in conventional boundary-layer interaction theory. A method is presented for computing the entrainment rate by an overlaid mixing layer model that accounts for the nonsimilar behavior and pressure gradients occurring in the near field region. An iterative scheme for coupling the model to analyses for the external inviscid flow, the external boundary layer, and the inviscid jet exhaust is also given. Results are presented that illustrate the qualitative behavior of the entrainment interaction under various flow conditions and that demonstrate the validity of the model by comparisons with experiment.

Wilmoth, R. G.↗

NASA NTRS1974

Measurement and correlation of aerodynamic heating to surface corrugation stiffened structures in thick turbulent boundary layers

The flow conditions for which heating distributions were measured on corrugated surfaces and wavy walls in turbulent boundary layers are shown, along with the ratio of the displacement thickness to the roughness height versus the local edge Mach number for an equivalent smooth surface. The present data are seen to greatly extend the range of data available on corrugated surfaces in turbulent boundary layers. These data were obtained by testing fullscale corrugation roughened panels in the wall boundary layer of a supersonic and hypersonic wind tunnel. The experimental program used to obtain the data is described. The data are analyzed and correlated in terms of the pertinent flow and geometric parameters. The developed correlations are compared with the available thin boundary layer data, as well as with previously published correlation techniques.

Brandon, H. J.↗

NASA NTRS1978

An improved analytical model of the separation region on boattail nozzles at subsonic speeds

A practical engineering calculation was developed to model the viscous effects of a separated, reverse flow region on afterbody pressures and drag. This viscous calculation was iteratively coupled with an inviscid flow calculation by means of an aerodynamic interface. A standard boundary layer displacement thickness was used to modify the afterbody shape where the flow was attached. A discriminating streamline calculation was developed to account for displacement effects of the reverse flow in separated regions with and without a flowing jet. The viscous flow calculation was coupled with a potential flow calculation. The analysis accurately predicted afterbody pressures and drag with variations in Reynolds number, Mach number, and afterbody shape.

Presz, W. M., Jr.↗

NASA NTRS1989

A posteriori estimates for shear correction factors in multilayered composite cylinders

A predictor-corrector approach is presented for calculating the composite shear correction factors and analyzing multilayered composite cylinders. In the predictor phase a two-dimensional first-order shear deformation theory is used to predict the gross response characteristics of the cylinder (vibration frequencies, average through-the-thickness displacements, rotations, and transverse shear strain energy per unit area) as well as the in-plane strains and stresses in the thickness direction. The three-dimensional equilibrium equations and constitutive relations are then used to compute the transverse stresses and strains as well as the transverse shear factors, and they are also used to correct the predicted response quantities of the cylinder. For simply supported multilayered cylinders the response quantities obtained by using the proposed approach are shown to be in close agreement with three-dimensional elasticity solutions for a wide range of lamination and geometric parameters. Also, the potential of the proposed approach for use in conjunction with large-scale finite element models of composite cylinders is outlined.

Noor, Ahmed K.↗

NASA NTRS1972

Experimental study of wall boundary layer growth in the 10 deg half angle conical nozzle of a reflected shock tunnel

Calibration studies of shock-tunnel nozzle flow were made in both N2 and Ar for a reservoir temperature of 2000 deg K and reservoir pressures of 15, 85, and 130 atm. The results for both test gases showed that the boundary layer was turbulent and the growth nonlinear, with the thickness being greater in Ar than N2 at comparable reservoir conditions. The boundary-layer thickness decreased with increasing reservoir pressure, but only small differences occurred between the two largest reservoir pressures. Good correlations of the boundary layer and displacement thickness were obtained for both N2 and Ar in terms of Reynolds number based on a reference temperature.

Menees, G. P.↗

NASA NTRSDate not supplied

Broadband Predictions of Optimized Proprotors in Axial Forward Flight

A low-fidelity toolchain was used to predict broadband self-noise trends for three proprotors in axial flight: a baseline (C24ND) and two acoustically constrained proprotors (OPT-III and COPR-3). Rotor loads were predicted with the ANOPP Propeller Analysis System (PAS) and self-noise was predicted with the semiempirical method of Brooks, Pope, and Marcolini (BPM) implemented in the ANOPP2 Self-Noise Internal Functional Module (ASNIFM). Comparisons to experimental data revealed that trends for turbulent boundary layer trailing edge (TBL-TE) noise could be modeled across several flight conditions by increasing the boundary layer thicknesses via the trip setting. Since the Mach number range of the BPM method is exceeded in these predictions, a dependence of boundary layer displacement thickness on blade station Mach number was suggested as a possible reason for needing to model thicker boundary layers, suggesting that the TBL-TE model needs to be developed further. Bluntness vortex shedding noise (BVS) predictions required tuning the trailing edge thickness and trailing edge closure angle for each flight condition to match experimental trends, demonstrating that the BVS noise model is incomplete and that BVS noise may vary with the angle of attack. This study indicates that the BPM self-noise method needs to be improved, which will lead to more accurate broadband predictions.

Broadband self-noise↗

NASA NTRSDate not supplied

Broadband Predictions of Optimized Proprotors in Axial Forward Flight

Broadband self-noise↗

NASA NTRS1980

Wind tunnel investigation of the Titan Forward Skirt compartment vent from a free-stream Mach number of 0.80 to 1.96

A test was conducted to determine the flow characteristics of the Titan forward skirt compartment vent over a free stream Mach number range of 0.80 to 1.96. The vent was mounted in a flat plate and the plate was flush mounted to the tunnel side wall with coinciding center lines. Air was discharged from a duct, located on the tunnel side wall behind the plate, through a canted aft 30 deg honeycomb vent into the free stream. Data for the analysis of the Titan forward skirt compartment venting during ascent through the atmosphere are provided. Full scale simulated flight hardware, such as the honeycomb vent, duct corrugations and field joint ring were used. Boundary layer thicknesses were used to vary boundary height. The highest vent discharge coefficient for any given Mach number and vent pressure ratio generally occurred at the maximum displacement thickness. With no vent flow the static pressure in the vent region was generally less than the free stream static pressure. With vent flow, the static pressures upstream of the vent increased, and those downstream of the vent decreased.

Johns, A. L.↗

NASA NTRS1975

Correlation of boundary layer quantities for hypersonic laminar flows with zero pressure gradient for several gases

The laminar boundary layer has been theoretically studied for six gases for flows over cold walls with zero pressure gradient at Mach numbers between 5.5 and 12.5 to correlate boundary layer quantities for the various gases. The flow conditions considered correspond to those that can be generated in test facilities such as the shock tunnel and the expansion tube. Computed results obtained using real gas properties indicate that the Eckert number based on edge conditions serves to correlate the results in terms of the wall shear stress and enthalpy gradient, the Stanton number, and the momentum thickness for the various gases within plus or minus 10 per cent for Te = Tw and Te approximately 3Tw. Computed Reynolds analogy factors exhibit very good agreement with those predicted by the Colburn analogy. Velocity and displacement thicknesses correlate well with Eckert number for Te = Tw, but fail to correlate for Te approximately 3Tw. Differences in results are traced to property variations. Results show that the Eckert number is a significant correlating variable for the flows considered.

Cook, W. J.↗

NASA NTRS2002

The Steady Flow Resistance of Perforated Sheet Materials in High Speed Grazing Flows

A study was conducted to determine the effects of high speed grazing air flow on the acoustic resistance of perforated sheet materials used in the construction of acoustically absorptive liners placed in commercial aircraft engine nacelles. Since DC flow resistance of porous sheet materials is known to be a major component of the acoustic resistance of sound suppression liners, the DC flow resistance of a set of perforated face-sheets and linear 'wiremesh' face-sheets was measured in a flow duct apparatus (up to Mach 0.8). Samples were fabricated to cover typical variations in perforated face-sheet parameters, such as hole diameter, porosity and sheet thickness, as well as those due to different manufacturing processes. The DC flow resistance data from perforated sheets were found to correlate strongly with the grazing flow Mach number and the face-sheet porosity. The data also show correlation against the boundary layer displacement thickness to hole-diameter ratio. The increase in resistance with grazing flow for punched aluminum sheets is in good agreement with published results up to Mach 0.4, but is significantly larger than expected above Mach 0.4. Finally, the tests demonstrated that there is a significant increase in the resistance of linear 'wiremesh' type face-sheet materials.

Syed, Asif A.↗

NASA NTRS1977

Experimental investigation of subsonic turbulent separated boundary layers on an airfoil

Detailed measurements of flow fields associated with low-speed turbulent boundary layers have been made for the 17% thick GA(W)-1 airfoil section at nominal angles of attack of 10, 14, and 18 deg; Reynolds number 2.2 million; and Mach number 0.13. The data include extensive pressure and velocity surveys of the pre- and post-separated regions on the airfoil and the associated wake. Integrated boundary-layer characteristics, including regions of separation on the airfoil, are also presented. The results indicate steep gradients of displacement thickness, momentum thickness, shape factor, and the separation streamline from the point of separation to the trailing edge of the airfoil. The present tests reveal that the region of flow reversal terminates within a suprisingly short distance of less than 20% chord downstream from the trailing edge for the test range of angle of attack.

Seetharam, H. C.↗