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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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NASA NTRS2000

MHD Energy Bypass Scramjet Performance with Real Gas Effects

The theoretical performance of a scramjet propulsion system incorporating an magneto-hydro-dynamic (MHD) energy bypass scheme is calculated. The one-dimensional analysis developed earlier, in which the theoretical performance is calculated neglecting skin friction and using a sudden-freezing approximation for the nozzle flow, is modified to incorporate the method of Van Driest for turbulent skin friction and a finite-rate chemistry calculation in the nozzle. Unlike in the earlier design, in which four ramp compressions occurred in the pitch plane, in the present design the first two ramp compressions occur in the pitch plane and the next two compressions occur in the yaw plane. The results for the simplified design of a spaceliner show that (1) the present design produces higher specific impulses than the earlier design, (2) skin friction substantially reduces thrust and specific impulse, and (3) the specific impulse of the MHD-bypass system is still better than the non-MHD system and typical rocket over a narrow region of flight speeds and design parameters. Results suggest that the energy management with MHD principles offers the possibility of improving the performance of the scramjet. The technical issues needing further studies are identified.

Park, Chul↗
NASA NTRS1981

Simple approximations for the asymptotic description of the interaction between a normal shock wave and a turbulent boundary layer at transonic speeds

The asymptotic description of the interaction between a normal shock wave and a turbulent boundary layer is reviewed. The layers necessary in a rational analysis of the interaction are discussed with emphasis on the differences from an interaction with a laminar boundary layer, the uncoupling of solutions for the distribution of pressure and skin friction at the wall, and the role of the Reynolds shear stress in these solutions. The accuracy of asymptotic solutions in flows at Reynolds numbers of technical interest is discussed. Solutions for the distribution of pressure and skin friction at the wall and the shape of the shock are considered for the case where the flow is near separation. For the pressure and skin friction, it is possible to write two simplified partial solutions, one valid at the beginning of the interaction and one valid somewhat downstream of the shock wave. A solution composed of these two parts and a linear interpolation between them appears to give good comparison with experiment; one unknown constant, independent of the parameters of the interaction, must be found from experiment. The simplified relations are presented. Comparison of numerical computations with experimental data indicates a possible value for the constant and shows quite satisfactory results.

Adamson, T. C., Jr.↗
NASA NTRS1988

Unsteady three-dimensional flow separation

A concise mathematical framework is constructed to study the topology of steady 3-D separated flows of an incompressible, or a compressible viscous fluid. Flow separation is defined by the existence of a stream surface which intersects with the body surface. The line of separation is itself a skin-friction line. Flow separation is classified as being either regular or singular, depending respectively on whether the line of separation contains only a finite number of singular points or is a singular line of the skin-friction field. The special cases of 2-D and axisymmetric flow separation are shown to be of singular type. In regular separation it is shown that a line of separation originates from a saddle point of separation of the skin-friction field and ends at nodal points of separation. Unsteady flow separation is defined relative to a coordinate system fixed to the body surface. It is shown that separation of an unsteady 3-D incompressible viscous flow at time t, when viewed from such a frame of reference, is topologically the same as that of the fictitious steady flow obtained by freezing the unsteady flow at the instant t. Examples are given showing effects of various forms of flow unsteadiness on flow separation.

Hui, W. H.↗
NASA NTRS1985

Airfoil large-eddy breakup devices for turbulent drag reduction

Large-eddy breakup devices were tested in a turbulent boundary layer at a momentum thickness Reynolds number of 7400. The drag reducing characteristics of thin, tapered plates, NACA 0009, and NACA 4409 airfoil shapes were determined and compared with lower Reynolds number data obtained with thin elements. The results indicate that at the higher Reynolds numbers of the present tests the thin, tapered plates in a tandem arrangement produced large downstream skin-friction reductions (maximum 0/40 percent/), and gave net drag reductions at least as large as for lower Reynolds numbers. Tandem NACA 0009 airfoils, which were approximately 10 times thicker than the thin plates, gave similar results with a maximum downstream skin-friction reduction of 0(30 percent) and a net drag reduction of 7 percent. The device drag for this shape was determined to be near laminar skin friction. The NACA 4409 airfoils suffered large regions of separated flow and produced no net drag reduction.

Anders, J. B.↗
NASA NTRS1979

An evaluation of linear acoustic theory for a hovering rotor

Linear acoustic calculations are compared with previously reported data for a small-scale hovering rotor operated at high tip Mach numbers. A detailed calculated description of the distributions of blade surface pressure and shear stress due to skin friction is presented. The noise due to skin friction and loading, in the rotor disk plane, is small compared to thickness noise. The basic conclusions of Boxwell et al about the importance of nonlinear effects are upheld. Some approximations involved in the current theories for the inclusion of nonlinear effects are discussed. Using a model nonlinear problem, it is shown that to use the acoustic analogy, good knowledge of the flowfield is required.

Morris, C. E. K., Jr.↗
NASA NTRS2010

Validation of Heat Transfer and Film Cooling Capabilities of the 3-D RANS Code TURBO

The capabilities of the 3-D unsteady RANS code TURBO have been extended to include heat transfer and film cooling applications. The results of simulations performed with the modified code are compared to experiment and to theory, where applicable. Wilcox s k-turbulence model has been implemented to close the RANS equations. Two simulations are conducted: (1) flow over a flat plate and (2) flow over an adiabatic flat plate cooled by one hole inclined at 35 to the free stream. For (1) agreement with theory is found to be excellent for heat transfer, represented by local Nusselt number, and quite good for momentum, as represented by the local skin friction coefficient. This report compares the local skin friction coefficients and Nusselt numbers on a flat plate obtained using Wilcox's k-model with the theory of Blasius. The study looks at laminar and turbulent flows over an adiabatic flat plate and over an isothermal flat plate for two different wall temperatures. It is shown that TURBO is able to accurately predict heat transfer on a flat plate. For (2) TURBO shows good qualitative agreement with film cooling experiments performed on a flat plate with one cooling hole. Quantitatively, film effectiveness is under predicted downstream of the hole.

Shyam, Vikram↗
NASA NTRS1987

An experimental study of the development of longitudinal vortex pairs embedded in a turbulent boundary layer

The mean streamwise development of pairs of longitudinal vortices embedded in an otherwise two-dimensional turbulent boundary layer was studied. Planes of closely spaced measurements of the three components of mean velocity were obtained at several streamwise locations, and the vorticity and circulation were calculated. Skin-friction measurements were also made. It was found that the rate of vorticity spreading in a vortex was greatly increased by close proximity of other vortices. The rate of streamwise circulation decrease was significantly greater for corotating vortices than for counter rotating vortices. Boundary-layer thinning and increased skin friction occured in regions where the secondary flow induced by the pairs was directed toward the wall; the boundary layer was thickened and skin friction reduced where the secondary flow was directed away from the wall.

Pauley, Wayne R.↗
NASA NTRS1976

Effect of geometry modifications on effectiveness of slot injection in hypersonic flow

The study presents wind-tunnel measurements of surface static pressures, equilibrium temperatures, and skin friction downstream of tangential slot injection into a thick turbulent hypersonic boundary layer from two modified slot configurations. The data are compared with results obtained for baseline configurations reported by Cary and Hefner (1970, 1972) to determine whether simple modifications to the slot configuration can produce improved cooling effectiveness and skin friction reduction. The baseline slot configurations are simply modified by thickening the slot lip and by elevating the location of the slot exit above the flat plate. Although the results indicate that simple modifications of the baseline slot configurations can enhance the skin friction reductions obtained with tangential slot injection, slot base drag estimates show that neither of the modifications will lessen the impact of the systems penalties for collecting, ducting, and injecting the slot air.

Hefner, J. N.↗
NASA NTRS1995

Venus' Free Obliquity

The predicted orientation of Venus' rotation axis relative to its orbit ran be uniquely determined given knowledge of its J2 gravity coefficient and polar moment of inertia C if its free obliquity is fully damped. This assumption seems warranted given the dominant damping mechanism: turbulent fluid friction at a core mantle boundary (CMB). This skin friction results from differential obliquity of mantle and core spin axes, and the associated damping rate could be as short as 1/10(exp 6) year. However, the observed pole orientation indicates a free obliquity amplitude epsilon approx. = 2.1 deg compared with a nominal forced amplitude of 0.5 deg. There are two plausible explanations. The most likely is that the observed obliquity is a tidally evolved end state in which core friction, modulated by CMB ellipticity and core obliquity amplitude, counterbalances solid and atmospheric tidal torques. This concept is similar to the explanation for the retrograde spin omega as an end state in which solid and atmospheric thermal tidal torques balance at the present spin rate because of the omega(sup -1) dependence of the axial thermal torque. Weaker effects such as solid and thermal tides can then compete with core friction and for plausible models, their sum tends to increase free obliquity. The obliquity balance is controlled by the nonlinear (and nearly quadratic) dependence of the CMB turbulent "skin friction" torque on obliquity.Core ellipticity also has a profound effect on tidal evolution of Venus obliquity.

Yoder, Charles F.↗
NASA NTRS1986

Combined Devices for Turbulent-Drag Reduction

Aircraft skin-friction drag reduced as much as 15 percent. One effective drag-reduction technique involves use of riblets. Riblets are longitudinal striations or grooves machined on originally smooth surface. Grooves alined with flow. Grooves have depths and spacings on order of turbulent wallstreak and burst dimensions and designed to change near-wall structure of turbulent boundary layer. Another approach, using large-eddy-breakup (LEBU) devices, or turbulence manipulators or ribbons also demonstrated reductions in local skin friction and net drag in air. LEBU device consists of thin, ribbonlike strips or airfoils suspended parallel to test surface and positioned within turbulent boundary layer. Technique potentially reduce net skin-friction drag by at least 15 percent on turbulent boundary layer of aircraft, representing possible annual savings in fuel costs of $300 to $400 million for U.S. commercial fleet. Also applicable to frictionloss reduction inside pipes and ducts, contributing to increased efficiency of pumps, heat exchangers, air conditioners, and other devices involving fluid flow.

Walsh, M. J.↗
NASA NTRS2018

Numerical Study of Turbulent Separation Bubbles with Varying Pressure Gradient and Reynolds Number

A family of cases each containing a small separation bubble is treated by direct numerical simulation (DNS), varying two parameters: the severity of the pressure gradients, generated by suction and blowing across the opposite boundary, and the Reynolds number. Each flow contains a well-developed entry region with essentially zero pressure gradient, and all are adjusted to have the same value for the momentum thickness, extrapolated from the entry region to the centre of the separation bubble. Combined with fully defined boundary conditions this will make comparisons with other simulations and turbulence models rigorous; we present results for a set of eight Reynolds-averaged Navier–Stokes turbulence models. Even though the largest Reynolds number is approximately 5.5 times higher than in a similar DNS study we presented in 1997, the models have difficulties matching the DNS skin friction very closely even in the zero pressure gradient, which complicates their assessment. In the rest of the domain, the separation location per se is not particularly difficult to predict, and the most definite disagreement between DNS and models is near reattachment. Curiously, the better models tend to cluster together in their predictions of pressure and skin friction even when they deviate from the DNS, although their eddy-viscosity levels are widely different in the outer region near the bubble (or they do not rely on an eddy viscosity). Stratford’s square-root law is satisfied by the velocity profiles, both at separation and reattachment. The Reynolds-number range covers a factor of two, with the Reynolds number based on the extrapolated momentum thickness equal to approximately 1500 and 3000. This allows tentative estimates of the improvements that even higher values will bring to the model comparisons. The solutions are used to assess models through pressure, skin friction and other measures; the flow fields are also used to produce effective eddy-viscosity targets for the models, thus guiding turbulence-modelling work in each region of the flow.

Coleman, G. N.↗
NASA NTRS1983

A study of the effects of Lebu devices on turbulent boundary layer drag

Initial measurements of the changes in local skin friction, velocity profile shape, and turbulence structure which result from the placement of tandem plates parallel to the wall in the outer region of thick turbulent boundary layers were made. Using a tunnel with a .75 m x 1.2 m x 7.3 m test section, which diverged so as to keep the pressure gradient less than 2x1000/ft, on the test wall, a skin friction reduction of approximately 30% was measured at xi/h = 62. This relaxed to a reduction of approximately 16% at xi/h = 124 for h/delta M = .6. The c sub f measurements for both the normal and modified boundary layers were obtained by measuring the slope of the velocity profile within the linear sublayer. Visual results indicated a continued presence of strong large eddy structure downstream of the devices. Local skin friction reduction of 12% at xi/h = 62 was also obtained with the manipulators above the boundary layer at y/delta m = 1.1.

Falco, R. E.↗
NASA NTRS1988

Correlations of velocity and temperature fluctuations in the stagnation-point flow of circular cylinder in turbulent flow

Boundary layer flow and turbulence transport analyses to study the influence of the free-stream turbulence on the surface heat transfer rate and the skin friction around the stagnation point of a circular cylinder in a turbulent flow are presented. The analyses are formulated with the turbulent boundary layer equations, the Reynolds stress transport equations and the k - epsilon two-equation turbulence modeling. The analyses are used to calculate the time-averaged turbulence double correlations, the mean flow properties, the surface heat transfer rate and the skin friction with an isotropic turbulence in the freestream. The analytical results are described and compared with the existing experimental measurements. Depending on the free-stream turbulence properties, the turbulence kinetic energy can increase or decrease as the flow moves toward the surface. However, the turbulence kinetic energy induces large Reynolds normal stresses at the boundary layer edge. The Reynolds normal stresses change the boundary layer profiles of the time-averaged double correlations of the velocity and temperature fluctuations, the surface heat transfer rate and the skin friction. The free-stream turbulence dissipation rate can affect the stagnation-point heat transfer rate but the influence of the free-stream temperature fluctuation on the heat transfer rate is insignificant.

Wang, Chi R.↗
NASA NTRS1950

Approximate Turbulent Boundary-Layer Development in Plane Compressible Flow Along Thermally Insulated Surfaces with Application to Supersonic-Tunnel Contour Correction

Numerical solutions of the differential equation obtained from the momentum theorem for the development of a turbulent boundary layer along a thermally insulated surface in two-dimensional and in radial shock-free flow are presented in tabular form for a range of Mach numbers from 0.100 to 10. The solution can be used in a step-wise procedure with any given distribution of favorable pressure gradients and for zero pressure gradients. Solutions are also given for use with moderate adverse pressure gradients. The mean velocity in the boundary layer is approximated by a power-law profile. In view of the stepwise integration methods to be used, the exponent designated the profile shape can be varied along the surface between the integral fraction limits 1/5 and 1/11 through interpolation. Agreement obtained between theoretical and experimental boundary-layer development in a supersonic nozzle at a nominal Mach number of 2 indicates the general validity of the approximations used in the analysis - in particular, the method of extrapolating low-speed skin-friction relations to high Mach number flows. The extrapolation method used assumes that the skin-friction coefficient depend primarily on Reynolds number, provided that the density and the kinematic viscosity are evaluated at surface conditions.

BOUNDARY LAYER - WIND TUNNEL WALLS↗
NASA NTRS1998

Prediction of Transitional Flows in the Low Pressure Turbine

Current turbulence models tend to give too early and too short a length of flow transition to turbulence, and hence fail to predict flow separation induced by the adverse pressure gradients and streamline flow curvatures. Our discussion will focus on the development and validation of transition models. The baseline data for model comparisons are the T3 series, which include a range of free-stream turbulence intensity and cover zero-pressure gradient to aft-loaded turbine pressure gradient flows. The method will be based on the conditioned N-S equations and a transport equation for the intermittency factor. First, several of the most popular 2-equation models in predicting flow transition are examined: k-e [Launder-Sharina], k-w [Wilcox], Lien-Leschiziner and SST [Menter] models. All models fail to predict the onset and the length of transition, even for the simplest flat plate with zero-pressure gradient(T3A). Although the predicted onset position of transition can be varied by providing different inlet turbulent energy dissipation rates, the appropriate inlet conditions for turbulence quantities should be adjusted to match the decay of the free-stream turbulence. Arguably, one may adjust the low-Reynolds-number part of the model to predict transition. This approach has so far not been very successful. However, we have found that the low-Reynolds-number model of Launder and Sharma [1974], which is an improved version of Jones and Launder [1972] gave the best overall performance. The Launder and Sharma model was designed to capture flow re-laminarization (a reverse of flow transition), but tends to give rise to a too early and too fast transition in comparison with the physical transition. The three test cases were for flows with zero pressure gradient but with different free-stream turbulent intensities. The same can be said about the model when considering flows subject to pressure gradient(T3C1). To capture the effects of transition using existing turbulence models, one approach is to make use of the concept of the intermittency to predict the flow transition. It was originally based on the intermittency distribution of Narasimha [1957], and then gradually evolved into a transport equation for the intermittency factor. Gostelow and associates [1994,1995] have made some improvements to Narasimha's method in an attempt to account for both favorable and adverse pressure gradients. Their approach is based on a linear, explicit combination of laminar and turbulent solutions. This approach fails to predict the overshoot of the skin friction on a flat plate near the end of transition zone, even though the length of transition is well predicted. The major flaw of Gostelow's approach is that it assumes the non-turbulent part being the laminar solution and the turbulent part being the turbulent solution and they do not interact across the transitional region. The technique in condition averaging the flow equations in intermittent flows was first introduced by Libby [1975] and Dopazo [1977] and further refined by Dick and associates [1988, 1996]. This approach employs two sets of transport equations for the non-turbulent part and the other for the turbulent part. The advantage of this approach is that it allows the interaction of non-turbulent and turbulent velocities through the introduction of additional source terms in the continuity and momentum equations for the non-turbulent and turbulent velocities. However, the strong coupling of the two sets of equations has caused some numerical difficulties, which requires special attention. The prediction of the skin friction can be improved by this approach via the implicit coupling of non-turbulent and turbulent velocity flelds. Another improvement of the interrmittency model can be further made by allowing the intermittency to vary in the cross-stream direction. This is one step prior to testing any proposal for the transport equation for the intermittency factor. Instead of solving the transport equation for the intermittency factor, the distribution for the intermittency factor is prescribed by Klebanoff's empirical formula [1955]. The skin friction is very well predicted by this new modification, including the overshoot of the profile near the end of the transition zone. The outcome of this study is very encouraging since it indicates that the proper description of the intermittency distribution is the key to the success of the model prediction. This study will be used to guide us on the modelling of the intermittency transport equation.

Huang, George↗
NASA NTRS1999

Use of Boundary Layer Transition Detection to Validate Full-Scale Flight Performance Predictions

Full-scale flight performance predictions can be made using CFD or a combination of CFD and analytical skin-friction predictions. However, no matter what method is used to obtain full-scale flight performance predictions knowledge of the boundary layer state is critical. The implementation of CFD codes solving the Navier-Stokes equations to obtain these predictions is still a time consuming, expensive process. In addition, to ultimately obtain accurate performance predictions the transition location must be fixed in the CFD model. An example, using the M2.4-7A geometry, of the change in Navier-Stokes solution with changes in transition and in turbulence model will be shown. Oil flow visualization using the M2.4-7A 4.0% scale model in the 14'x22' wind tunnel shows that fixing transition at 10% x/c in the CFD model best captures the flow physics of the wing flow field. A less costly method of obtaining full-scale performance predictions is the use of non-linear Euler codes or linear CFD codes, such as panel methods, combined with analytical skin-friction predictions. Again, knowledge of the boundary layer state is critical to the accurate determination of full-scale flight performance. Boundary layer transition detection has been performed at 0.3 and 0.9 Mach numbers over an extensive Reynolds number range using the 2.2% scale Reference H model in the NTF. A temperature sensitive paint system was used to determine the boundary layer state for these conditions. Data was obtained for three configurations: the baseline, undeflected flaps configuration; the transonic cruise configuration; and, the high-lift configuration. It was determined that at low Reynolds number conditions, in the 8 to 10 million Reynolds number range, the baseline configuration has extensive regions of laminar flow, in fact significantly more than analytical skin-friction methods predict. This configuration is fully turbulent at about 30 million Reynolds number for both 0.3 and 0.9, Mach numbers. Both the transonic cruise and the high-lift configurations were fully turbulent aft of the leading-edge flap hingeline at all Reynolds numbers.

Hamner, Marvine↗
NASA NTRS1985

The effects of compliant walls on transition and turbulence

Two theoretical approaches to the phenomenon of drag reduction by a compliant wall are proposed. One is to study the limit cycle oscillations of the Navier-Stokes equations for a boundary layer flow over a compliant wall or a Poiseuille flow in a channel with a compliant wall. The other is a complementary study of a simpler, sequential model, which partially decouples the fluid and wall motions. In this approach, a 'Total Model' is constructed from suitable component models. While the first study is more rigorously based, it will be initially limited, by computational constraints, to relatively low Reynolds numbers. The second, more approximate approach, will allow consideration of a fully developed turbulent flow. The Total Model will predict the compliant wall motion accurately, but will not per se be sufficient to predict the skin friction drag. However, once the wall motion is known, the skin friction drag and the pressure drag can be computed separately, using one of the existing models.

Dowell, E. H.↗