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At least 307 records · Page 17

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NASA NTRS1990

The stability of a three dimensional laminar boundary layer over a swept flat plate

The linear stability of the laminar boundary layer on a swept flat plate with an imposed favorable pressure gradient was studied utilizing a linear stability model which accounts for streamline curvature for three-dimensional incompressible flows. Calculations were performed for a leading-edge sweep angle of 45 degrees and freestream valocity of 19 m/s. Computed disturbance amplification rates and wavelengths for stationary crossflow vortices were compared with available experimental results.

Mueller, B.↗

NASA NTRS1940

Tests of a Gust-alleviating Flap in the Gust Tunnel

Tests were made to determine the effectiveness of a long-period dynamically overbalanced flap in reducing airplane accelerations due to atmospheric gusts. For two gust shapes, one gust velocity, one forward velocity, and one wing loading, a series of flights was made with the flap locked and was then repeated with the flap free to operate. The records were evaluated by routine methods. The results indicate that the flap reduced the maximum acceleration increment 39 percent for a severe gust but with a representative gust shape (a sharp-edge gust), the reduction was only 3 percent. The results also indicate that the flap tended to reduce the longitudinal stability of the airplane. Computations made of the effectiveness and the action of the flap were in good agreement with the experimental results.

Donely, Philip↗

NASA NTRS1960

Effects of Sweep Angle on the Boundary-Layer Stability Characteristics of an Untapered Wing at Low Speeds

An investigation was conducted in the Ames 12-Foot Low-Turbulence Pressure Tunnel to determine the effects of sweep on the boundary-layer stability characteristics of an untapered variable-sweep wing having an NACA 64(2)A015 section normal to the leading edge. Pressure distribution and transition were measured on the wing at low speeds at sweep angles of 0, 10, 20, 30, 40, and 50 deg. and at angles of attack from -3 to 3 deg. The investigation also included flow-visualization studies on the surface at sweep angles from 0 to 50 deg. and total pressure surveys in the boundary layer at a sweep angle of 30 deg. for angles of attack from -12 to 0 deg. It was found that sweep caused premature transition on the wing under certain conditions. This effect resulted from the formation of vortices in the boundary layer when a critical combination of sweep angle, pressure gradient, and stream Reynolds number was attained. A useful parameter in indicating the combined effect of these flow variables on vortex formation and on beginning transition is the crossflow Reynolds number. The critical values of crossflow Reynolds number for vortex formation found in this investigation range from about 135 to 190 and are in good agreement with those reported in previous investigations. The values of crossflow Reynolds number for beginning transitions were found to be between 190 and 260. For each condition (i.e., development of vortices and initiation of transition at a given location) the lower values in the specified ranges were obtained with a light coating of flow-visualization material on the surface. A method is presented for the rapid computation of crossflow Reynolds number on any swept surface for which the pressure distribution is known. From calculations based on this method, it was found that the maximum values of crossflow Reynolds number are attained under conditions of a strong pressure gradient and at a sweep angle of about 50 deg. Due to the primary dependence on pressure gradient, effects of sweep in causing premature transition are generally first encountered on the lower surfaces of wings operating at positive angles of attack.

Boltz, Frederick W.↗

NASA NTRS1999

Step Bunching: Influence of Impurities and Solution Flow

Step bunching results in striations even at relatively early stages of its development and in inclusions of mother liquor at the later stages. Therefore, eliminating step bunching is crucial for high crystal perfection. At least 5 major effects causing and influencing step bunching are known: (1) Basic morphological instability of stepped interfaces. It is caused by concentration gradient in the solution normal to the face and by the redistribution of solute tangentially to the interface which redistribution enhances occasional perturbations in step density due to various types of noise; (2) Aggravation of the above basic instability by solution flowing tangentially to the face in the same directions as the steps or stabilization of equidistant step train if these flows are antiparallel; (3) Enhanced bunching at supersaturation where step velocity v increases with relative supersaturation s much faster than linear. This v(s) dependence is believed to be associated with impurities. The impurities of which adsorption time is comparable with the time needed to deposit one lattice layer may also be responsible for bunching; (4) Very intensive solution flow stabilizes growing interface even at parallel solution and step flows; (5) Macrosteps were observed to nucleate at crystal corners and edges. Numerical simulation, assuming step-step interactions via surface diffusion also show that step bunching may be induced by random step nucleation at the facet edge and by discontinuity in the step density (a ridge) somewhere in the middle of a face. The corresponding bunching patterns produce the ones observed in experiment. The nature of step bunching generated at the corners and edges and by dislocation step sources, as well as the also relative importance and interrelations between mechanisms 1-5 is not clear, both from experimental and theoretical standpoints. Furthermore, several laws controlling the evolution of existing step bunches have been suggested, though unambiguous conclusions are still missing. Addressing these issues is the major goal of the present project. The theory addressing the above problem, experimental methods, several figures which include: (1) the spatial wave numbers at which the system is neutrally stable as a function of growth velocity for linear kinetics and supersaturation for nonlinear kinetics; (2) a schematic of the experiment of lysozyme crystal growing under conditions of natural convection; (3) fluctuations in time, t, of the normal growth rate, R(t), vicinal slope, p(t) and Fourier Spectra of R(t), discussions and conclusions are presented.

Chernov, A. A.↗

NASA NTRS2008

Effects of Wall Cooling on Hypersonic Boundary Layer Receptivity Over a Cone

Effects of wall cooling on the receptivity process induced by the interaction of slow acoustic disturbances in the free-stream are numerically investigated for a boundary layer flow over a 5-degrees straight cone. The free-stream Mach number is 6.0 and the Reynolds number is 7.8x10(exp 6)/ft. Both the steady and unsteady solutions are obtained by solving the full Navier-Stokes equations using 5th-order accurate weighted essentially non-oscillatory (WENO) scheme for space discretization and using 3rd-order total variation diminishing (T VD) Runge-K utta scheme for time integration. Computations are performed for a cone with nose radius of 0.001 inch for adiabatic wall temperature (T(sub aw)), 0.75*T(sub aw), 0.5*T(sub aw), 0.40*T(sub aw), 0.30*T(sub aw), and 0.20*T(sub aw). Once the mean flow field is computed, disturbances are introduced at the upstream end of the computational domain. Generation of instability waves from leading edge region and receptivity of boundary layer to slow acoustic waves are investigated. Computations showed that wall cooling has strong stabilization effect on the first mode disturbances as was observed in the experiments. T ransition location moved to upstream when wall cooling was applied It is also found that the boundary layer is much more receptive to fast acoustic wave (by almost a factor of 50). When simulations performed using the same forcing frequency growth of the second mode disturbances are delayed with wall cooling and they attained values two times higher than that of adiabatic case. In 0.20*T(sub aw) case the transition Reynolds number is doubled compared to adiabatic conditions. The receptivity coefficient for adiabatic wall case (804 R) is 1.5225 and for highly cooled cones (241, and 161 R); they are in the order of 10(exp -3).

Kara, K.↗

NASA NTRS1953

Low-Speed Longitudinal Stability Characteristics of a 1/6-Scale Model of the Republic XF-84H Airplane with the Propeller Operating

A low-speed investigation was made of a 1/6-scale model of the Republic XF-84H airplane. The model had a single tractor propeller and a 40deg swept wing of aspect ratio 3.45. This investigation was undertaken to provide information on the effects of propeller operation on longitudinal stability characteristics for the XF -84H airplane and to provide an indication of slipstream effects that might be encountered on similar swept-wing configurations. Effects of propeller operation were generally destabilizing for all conditions investigated; however, the over-all stability characteristics with power on were greatly dependent on the power-off characteristics. With flaps and slats retracted, longitudinal instability was present at moderate angles of attack both with the propeller off and with power on. The longitudinal stability with flaps and slats deflected, which was satisfactory without power, was decreased by propeller operation, but no marked pitch-up tendency was indicated. Significant improvement in the power-on stability with flaps retracted was achieved by use of either a wing fence at 75 percent semispan, a leading-edge chord-extension from 65 to 94 percent semispan, or a raised horizontal tail located 65 percent semispan above the thrust line.

Sleeman, William C.↗

NASA NTRS2013

Receptivity of Hypersonic Boundary Layers to Distributed Roughness and Acoustic Disturbances

Boundary-layer receptivity and stability of Mach 6 flows over smooth and rough seven-degree half-angle sharp-tipped cones are numerically investigated. The receptivity of the boundary layer to slow acoustic disturbances, fast acoustic disturbances, and vortical disturbances is considered. The effects of three-dimensional isolated roughness on the receptivity and stability are also simulated. The results for the smooth cone show that the instability waves are generated in the leading edge region and that the boundary layer is much more receptive to slow acoustic waves than to the fast acoustic waves. Vortical disturbances also generate unstable second modes, however the receptivity coefficients are smaller than that of the slow acoustic wave. Distributed roughness elements located near the nose region decreased the receptivity of the second mode generated by the slow acoustic wave by a small amount. Roughness elements distributed across the continuous spectrum increased the receptivity of the second mode generated by the slow and fast acoustic waves and the vorticity wave. The largest increase occurred for the vorticity wave. Roughness elements distributed across the synchronization point did not change the receptivity of the second modes generated by the acoustic waves. The receptivity of the second mode generated by the vorticity wave increased in this case, but the increase is lower than that occurred with the roughness elements located across the continuous spectrum. The simulations with an isolated roughness element showed that the second mode waves generated by the acoustic disturbances are not influenced by the small roughness element. Due to the interaction, a three-dimensional wave is generated. However, the amplitude is orders of magnitude smaller than the two-dimensional wave.

Balakumar, P.↗

NASA NTRS1981

SPF/DB primary structure for supersonic aircraft (T-38 horizontal stabilizer)

The structural integrity and potential cost savings of superplastic forming/diffusion bonding (SPF/DB) titanium structure for future Supersonic Cruise Research (SCR) and military aircraft primary structure applications was demonstrated. Using the horizontal stabilizer of the T-38 aircraft as a baseline, the structure was redesigned to the existing criteria and loads, using SPF/DB titanium technology. The general concept of using a full-depth sandwich structure which is attached to a steel spindle, was retained. Trade studies demonstrated that the optimum design should employ double-truss, sinewave core in the deepest section of the surface, making a transition to single-truss core in the thinner areas at the leading and trailing edges and at the tip. At the extreme thin edges of the surface, the single-truss core was changed to dot core to provide for gas passages during the SPF/DB process. The selected SPF/DB horizontal stabilizer design consisted of a one-piece SPF/DB sinewave truss core panel, a trunnion fitting, and reinforcing straps. The fitting and the straps were mechanically fastened to the SPF/DB panel.

Delmundo, A. R.↗

NASA NTRS1951

Effect of Ground Interference on the Aerodynamic Characteristics of a 42 Degrees Sweptback Wing

The effects of ground interference on the aerodynamic characteristics of a 42 degrees sweptback wing have been determined at distances above the ground 0.68 and 0.92 of the mean aerodynamic chord (measured from the 0.25 mean aerodynamic chord). The wing was tested without flaps and with inboard trailing-edge split and outboard leading-edge flaps deflected. The wing had an aspect ratio of 4, a taper ratio of 0.625, and NACA 641-112 airfoil sections perpendicular to the 0.273 chord line. The results are, in general, comparable to those reported for unswept wings. The longitudinal stability at the stall was not materially affected at the ground heights of the present tests.

RESEARCH TECHNIQUE, AERODYNAMICS↗

NASA NTRS2023

Fully Coupled Aeroelastic Stability Analysis of Adaptive Shape Memory Alloy Structural Technologies for Airframe Noise Reduction

This final report documents work performed by ATA Engineering, Inc., (ATA) to develop computational models and analyze the coupled fluid-structure response of two types of noise treatments applied to the leading-edge-slat component of a high-lift system typical of modern transport aircraft. The first treatment is a slat-gap filler (SGF), which closes the gap between the suction surfaces of a deployed slat and an aircraft main wing, and the second treatment is a slat-cove filler (SCF), which replaces the recirculating flow on the slat cove with a surface that promotes flow attachment. The representative airframe chosen for this work was NASA’s High-Lift Common Research Model (CRM-HL) in a baseline high-lift configuration. Superelastic shape memory alloys (SMAs) have been identified as enabling materials for these structural treatments. Since the technology elements rely upon having a highly reconfigurable structure, designs were assessed for their static aeroelastic deflection as well as their dynamic aeroelastic stability using coupled computational fluid dynamics (CFD) and nonlinear computational structural dynamics (NL-CSD) tools. Specifically, fluid-structure interaction (FSI) problems were solved computationally using the CFD solver Loci/CHEM and the NL-CSD solver Abaqus. As a part of the overall project, a similar capability was implemented using the CFD solver FUN3D coupled to Abaqus, although that work is documented in a separate report and that FSI framework was not used to analyze any of NASA’s SGF and SCF configurations. The technical approach consisted of solving for the flow field around the entire vehicle using a global CFD model, followed by extraction of relevant local subdomain data for CFD and NL-CSD co-simulations. The SGF design was analyzed using both 2D and 3D co-simulations to predict quasi-static aeroelastic deformations and to assess dynamic aeroelastic stability, whereas the SCF was analyzed in 2D only. SGF static aeroelastic response predictions focused on characterizing the deformed shape, with maximum displacements predicted to be on the order of magnitude of the technology element panel thickness. SGF dynamic aeroelastic response predictions used Partial Floquet analysis of the temporal evolution of selected nodal displacements to quantify the sign and magnitude of aeroelastic damping. Results suggest that the CRM-HL operating conditions would result in a dynamically stable response. The simulated dynamic pressure was also increased up to a factor of about four, and resulting responses suggest that predicted dynamic stability would be achieved with some margin.

Fluid Structure Interaction↗

NASA NTRS2014

In-Flight Aeroelastic Stability of the Thermal Protection System on the NASA HIAD, Part I: Linear Theory

Conical shell theory and piston theory aerodynamics are used to study the aeroelastic stability of the thermal protection system (TPS) on the NASA Hypersonic Inflatable Aerodynamic Decelerator (HIAD). Structural models of the TPS consist of single or multiple orthotropic conical shell systems resting on several circumferential linear elastic supports. The shells in each model may have pinned (simply-supported) or elastically-supported edges. The Lagrangian is formulated in terms of the generalized coordinates for all displacements and the Rayleigh-Ritz method is used to derive the equations of motion. The natural modes of vibration and aeroelastic stability boundaries are found by calculating the eigenvalues and eigenvectors of a large coefficient matrix. When the in-flight configuration of the TPS is approximated as a single shell without elastic supports, asymmetric flutter in many circumferential waves is observed. When the elastic supports are included, the shell flutters symmetrically in zero circumferential waves. Structural damping is found to be important in this case. Aeroelastic models that consider the individual TPS layers as separate shells tend to flutter asymmetrically at high dynamic pressures relative to the single shell models. Several parameter studies also examine the effects of tension, orthotropicity, and elastic support stiffness.

Goldman, Benjamin D.↗

NASA NTRS1977

Aerodynamic characteristics of an F-8 aircraft configuration with a variable camber wing at Mach numbers from 1.5 to 2.0

A 0.1-scale model of an F-8 aircraft was tested over a range of Mach numbers from 1.5 to 2.0. Reynolds number of 4.12 million was based on wing mean-aerodynamic chord for angles of attack varying from -2 deg to +12 deg. The model was equipped with an advanced-technology-conformal-variable-camber wing (ATCVCW) having simple hinge flaps. Data were also obtained for the model equipped with the basic F-8 wing and conventional flaps. Model variables included aileron and wing trailing edge deflections and horizontal tail incidence. The ATCVCW configuration produced slight improvements in lift-curve slope, drag, and static longitudinal stability over that of the basic F-8 wing configuration. Flap effectiveness was essentially the same for both wings.

Boltz, F. W.↗

NASA NTRS1983

Aerodynamic characteristics of generalized bent biconic bodies for aero-assisted, orbital-transfer vehicles

A method was developed to generate the surface coordinates of body shapes suitable for aeroassisted, orbital-transfer vehicles (AOTVs) by extending bent biconic geometries. Lift, drag, and longitudinal moments were calculated for the bodies using Newtonian flow theory. These techniques were applied to symmetric and asymmetric aerobraking vehicles, and to an aeromaneuvering vehicle with high L/D. Results for aerobraking applications indicate that a 70 deg, fore half cone angle with a spherically blunted nose, rounded edges, and a slight asymmetry would be appropriate. Moreover, results show that an aeromaneuvering vehicle with L/D 2.0, and with sufficient stability, is feasible.

Davies, C. B.↗

NASA NTRS1983

The stability of coronal loops - Finite-length and pressure-profile limits

Results are described from a quickly converging, necessary-and-sufficient, MHD-stability test for coronal-loop models. The primary stabilizing influence arises from magnetic line tying at the photosphere, and this end conditions requires a series expansion of possible loop excitations. The stability boundary is shown to quickly approach a limit as the number of terms increases, providing a critical length for the loop in proportion to its transverse magnetic scale. Several models of force-free-field profiles are tested and the stability behavior of a localized current channel, embedded in an external current-free region, is shown to be superior to that of other, broader, current profiles. Pressure-gradient effects, leading to increased or decreased stability, are shown to be amplified by line tying. Long loops must either conduct low net current, or exhibit an axial-field reversal coexisting with a low-pressure core. The limits on stability depend on the magnetic aspect ratio, the plasma-to-magnetic pressure ratio, and the field orientation at the loop edge. Applications of these results to the structure of coronal loops are described.

Einaudi, G.↗

NASA NTRS1985

Long-life high performance fuel cell program

A multihundred kilowatt Regenerative Fuel Cell for use in a space station is envisioned. Three 0.508 sq ft (471.9 cm) active area multicell stacks were assembled and endurance tested. The long term performance stability of the platinum on carbon catalyst configuration suitability of the lightweight graphite electrolyte reservoir plate, the stability of the free standing butyl bonded potassium titanate matrix structure, and the long life potential of a hybrid polysulfone cell edge frame construction were demonstrated. A 18,000 hour demonstration test of multicell stack to a continuous cyclical load profile was conducted. A total of 12,000 cycles was completed, confirming the ability of the alkaline fuel cell to operate to a load profile simulating Regenerative Fuel Cell operation. An orbiter production hydrogen recirculation pump employed in support of the cyclical load profile test completed 13,000 hours of maintenance free operation. Laboratory endurance tests demonstrated the suitability of the butyl bonded potassium matrix, perforated nickel foil electrode substrates, and carbon ribbed substrate anode for use in the alkaline fuel cell. Corrosion testing of materials at 250 F (121.1 C) in 42% wgt. potassium identified ceria, zirconia, strontium titanate, strontium zirconate and lithium cobaltate as candidate matrix materials.

Martin, R. E.↗

NASA NTRS1990

The stability of a three-dimensional laminar boundary layer on a swept flat plate

The linear stability of the laminar boundary layer on a swept flat plate with an imposed favorable pressure gradient was studied utilizing a linear stability model which accounts for streamline curvature for three-dimensional incompressible flows. Calculations were performed for an effective leading-edge sweep angle of 42.5 degrees and freestream velocity of 19 m/s. Computed disturbance amplification rates for the spectrum of amplified frequencies and wavelengths for stationary crossflow vortices were compared with experimental results.

Collier, F. S., Jr.↗

NASA NTRS1955

The Compressible Laminar Boundary Layer with Fluid Injection

A solution of the equations of the compressible laminar boundary layer including the effects of transpiration cooling is presented. The analysis applies to the flow over an isothermal porous plate with a velocity of fluid injection proportional to the reciprocal of the square root of the distance from the leading edge. The effect of several flow parameters on coolant-flow rates is discussed with the aid of representative examples. A stability analysis indicates that, although transpiration cooling requires a lower surface temperature for stable flow than does internal wall cooling, this lower temperature can be obtained with a smaller expenditure of coolant.

Low, George M↗

NASA NTRS1995

A parametric study of supersonic laminar flow for swept wings using linear stability analysis

A parametric study to predict the extent of laminar flow on the upper surface of a generic swept-back wing (NACA 64A010 airfoil section) at supersonic speeds was conducted. The results were obtained by using surface pressure predictions from an Euler/Navier-Stokes computational fluid dynamics code coupled with a boundary layer code, which predicts detailed boundary layer profiles, and finally with a linear stability code to determine the extent of laminar flow. The parameters addressed are Reynolds number, angle of attack, and leading-edge wing sweep. The results of this study show that an increase in angle of attack, for specific Reynolds numbers, can actually delay transition. Therefore, higher lift capability, caused by the increased angle of attack, as well as a reduction in viscous drag due to the delay in transition is possible for certain flight conditions.

Cummings, Russell M.↗