Kmak, Frank J.
The Unitary Plan Wind Tunnel (UPWT) was modernized to improve performance, capability, productivity, and reliability. Automation systems were installed in all three UPWT tunnel legs and the Auxiliaries facility. Major improvements were made to the four control rooms, model support systems, main drive motors, and main drive speed control. Pressure vessel repairs and refurbishment to the electrical distribution system were also completed. Significant changes were made to improve test section flow quality in the 11-by 11-Foot Transonic leg. After the completion of the construction phase of the project, acceptance and checkout testing was performed to demonstrate the capabilities of the modernized facility. A pneumatic test of the tunnel circuit was performed to verify the structural integrity of the pressure vessel before wind-on operations. Test section turbulence, flow angularity, and acoustic parameters were measured throughout the tunnel envelope to determine the effects of the tunnel flow quality improvements. The new control system processes were thoroughly checked during wind-off and wind-on operations. Manual subsystem modes and automated supervisory modes of tunnel operation were validated. The aerodynamic and structural performance of both the new composite compressor rotor blades and the old aluminum rotor blades was measured. The entire subsonic and supersonic envelope of the 11-by 11-Foot Transonic leg was defined up to the maximum total pressure.
Kmak, Frank J.; Hudgins, M.; Hergert, D.; George, Michael W. (Technical Monitor)
The 11-By 11-Foot Transonic leg of the Unitary Plan Wind Tunnel (UPWT) was modernized to improve tunnel performance, capability, productivity, and reliability. Wind tunnel tests to demonstrate the readiness of the tunnel for a return to production operations included an Integrated Systems Test (IST), calibration tests, and airplane validation tests. One of the two validation tests was a 0.037-scale Boeing 777 model that was previously tested in the 11-By 11-Foot tunnel in 1991. The objective of the validation tests was to compare pre-modernization and post-modernization results from the same airplane model in order to substantiate the operational readiness of the facility. Evaluation of within-test, test-to-test, and tunnel-to-tunnel data repeatability were made to study the effects of the tunnel modifications. Tunnel productivity was also evaluated to determine the readiness of the facility for production operations. The operation of the facility, including model installation, tunnel operations, and the performance of tunnel systems, was observed and facility deficiency findings generated. The data repeatability studies and tunnel-to-tunnel comparisons demonstrated outstanding data repeatability and a high overall level of data quality. Despite some operational and facility problems, the validation test was successful in demonstrating the readiness of the facility to perform production airplane wind tunnel%, tests.
Nguyen, Nhan; Ardema, Mark
This paper describes a modeling method and a new optimal control approach to investigate a Mach number control problem for the NASA Ames 11-Foot Transonic Wind Tunnel. The flow in the wind tunnel is modeled by the 1-D unsteady Euler equations whose boundary conditions prescribe a controlling action by a compressor. The boundary control inputs to the compressor are in turn controlled by a drive motor system and an inlet guide vane system whose dynamics are modeled by ordinary differential equations. The resulting Euler equations are thus coupled to the ordinary differential equations via the boundary conditions. Optimality conditions are established by an adjoint method and are used to develop a model predictive linear-quadratic optimal control for regulating the Mach number due to a test model disturbance during a continuous pitch
Ulbrich, Norbert; Boone, Alan R.
Data from the test of a large semispan model was used to perform a direct validation of a wall interference correction system for a transonic slotted wall wind tunnel. At first, different sets of uncorrected aerodynamic coefficients were generated by physically changing the boundary condition of the test section walls. Then, wall interference corrections were computed and applied to all data points. Finally, an interpolation of the corrected aerodynamic coefficients was performed. This interpolation made sure that the corrected Mach number of a given run would be constant. Overall, the agreement between corresponding interpolated lift, drag, and pitching moment coefficient sets was very good. Buoyancy corrections were also investigated. These studies showed that the accuracy goal of one drag count may only be achieved if reliable estimates of the wall interference induced buoyancy correction are available during a test.
Schnell, W. C.; Ordonez, G. W.
A 1/8 scale jet-effects model was tested in the NASA Ames 11 ft transonic tunnel at static conditions and over a range of Mach numbers from 0.4 to 1.4. The data presented show that significant differences in aeropropulsion performance can be expected by varying the exhaust nozzle type and its geometric parameters on a V/STOL underwing nacelle installation.
Hawke, Veronica M.
The test section of the 11 by 11-foot Unitary Plan Transonic Wind Tunnel (11-foot UPWT) may receive an upgrade of larger optical windows on both the North and South sides. These new larger windows will provide better access for optical imaging of test article flow phenomena including surface and off body flow characteristics. The installation of these new larger windows will likely produce a change to the aerodynamic characteristics of the flow in the Test Section. In an effort understand the effect of this change, a computational model was employed to predict the flows through the slotted walls, in the test section and around the model before and after the tunnel modification. This report documents the solid CAD model that was created and the inviscid computational analysis that was completed as a preliminary estimate of the effect of the changes.
Dods, J. B., Jr.
The static and dynamic rotor blade stresses of the three stage compressor were measured. Data are presented in terms of total blade stress for the complete operational range of compressor speeds and tunnel total pressures. Modal frequencies and variations with tunnel conditions were measured. Phase angles and coherences between various gage combinations are also presented. Recommendations for improvements are given for future rotor blade experimental investigations.
Nichols, M. E.
The results are documented of jet plume effects wind tunnel test of the 0.020-scale 88-OTS launch configuration space shuttle vehicle model in the 11 x 11 foot leg of the NASA/Ames Research Center Unitary Plan Wind Tunnel. This test involved cold gas main propulsion system (MPS) and solid rocket motor (SRB) plume simulations at Mach numbers from 0.6 to 1.4. Integrated vehicle surface pressure distributions, elevon and rudder hinge moments, and wing and vertical tail root bending and torsional moments due to MPS and SRB plume interactions were determined. Nozzle power conditions were controlled per pretest nozzle calibrations. Model angle of attack was varied from -4 deg to +4 deg; model angle of sideslip was varied from -4 deg to +4 deg. Reynolds number was varied for certain test conditions and configurations, with the nominal freestream total pressure being 14.69 psia. Plotted force and pressure data are presented.
Results of a jet plume effects test on Rockwell International integrated space shuttle vehicle using a vehicle 5 configuration 0.02-scale model (88-OTS) in the 11 by 11 foot leg of the NASA/Ames Research Center unitary plan wind tunnel (IA19), volume 1
Nichols, M. E.
Results are presented of jet plume effects test IA19 using a vehicle 5 configuration integrated space shuttle vehicle 0.02-scale model in the NASA/Ames Research Center 11 x 11-foot leg of the unitary plan wind tunnel. The jet plume power effects on the integrated vehicle static pressure distribution were determined along with elevon, main propulsion system nozzle, and solid rocket booster nozzle effectiveness and elevon hinge moments.
Terminal area energy management regime investigations utilizing an 0.030-scale model (47-0) of the space shuttle vehicle orbiter configuration 140A/B/C/R in the Ames Research Center 11 x 11 foot transonic wind tunnel (OA148), volume 5
Hawthorne, P. J.
Data obtained in wind tunnel test OA148 are presented. The objectives of the test series were to: (1) obtain pressure distributions, forces and moments over the vehicle 5 orbiter in the thermal area energy management (TAEM) and approach phases of flight; (2) obtain elevon and rudder hinge moments in the TAEM and approach phases of flight; (3) obtain body flap and elevon loads for verification of loads balancing with integrated pressure distributions; and (4) obtain pressure distributions near the short OMS pods in the high subsonic, transonic and low supersonic Mach number regimes.
Terminal area energy management regime investigations utilizing an 0.030-scale model (47-0) of the space shuttle vehicle orbiter configuration 140A/B/C/R in the Ames Research Center 11 x 11 foot transonic wind tunnel (0A148), volume 1
Hawthorne, P. J.
Data obtained in wind tunnel tests are presented. The objectives of the tests were to: (1) obtain pressure distributions, forces and moments over the vehicle 5 Orbiter in the terminal area energy management (TAEM) and approach phases of flight; (2) obtain elevon and rudder hinge moments in the TAEM and approach phases of flight; (3) obtain body flap and elevon loads for verification of loads balancing with integrated pressure distributions; and (4) obtain pressure distributions near the short OMS pods in the high subsonic, transonic and low supersonic Mach number regimes. Testing was conducted over a Mach number range from 0.6 to 1.4 with Reynolds number variations from 4.57 million to 2.74 million per foot. Model angle-of-attack was varied from -4 to 16 degrees and angles of side slip ranged from -8 to 8 degrees.
Terminal area energy management regime investigations utilizing an 0.030-scale model (47-0) of the space shuttle vehicle orbiter configuration 140A/B/C/R in the Ames Research Center 11 x 11 foot transonic wind tunnel (OH/48)
Hawthorne, P. J.
Data obtained in a wind tunnel test were examined to: (1) obtain pressure distributions, forces and moments over the vehicle 5 Orbiter in the terminal area energy management (TAEM) and approach phases of flight; (2) obtain elevon and rudder hinge moments in the TAEM and approach phases of flight; (3) obtain body flap and elevon loads for verification of loads balancing with integrated pressure distributions; and (4) obtain pressure distributions near the short OMS pods in the high subsonic, transonic and low supersonic Mach number regimes. Testing was conducted over a Mach number range from 0.6 to 1.4 with Reynolds number variations from 7.57 x 1 million to 2.74 x 1 million per foot. Model angle of attack was varied from -4 to 16 degrees and angles of sideslip ranged from -8 to 8 degrees.
Sleppy, Mark A.; Engel, Eric A.; Watson, Kevin T.; Atler, Douglas M.
In an effort to achieve the maximum possible Reynolds number (Re) when conducting production testing for flight loads aerodynamic databases, it has been the preferred practice of The Boeing Company / Commercial Airplanes (BCA) -- Loads and Dynamics Group since the early 1990's to test large scale semi-span models in the 11- By 11-Foot Transonic Wind Tunnel (TWT) leg of the Unitary Plan Wind Tunnel (UPWT) at the NASA Ames Research Center (ARC). There are many problems related to testing large scale semi-span models of high aspect ratio flexible transport wings, such as; floor boundary layer effects, wing spanwise wall effects, solid blockage buoyancy effects, floor mechanical interference effects, airflow under the model effects, or tunnel flow gradient effects. For most of these issues, BCA has developed and implemented either standard testing methods or numerical correction schemes and these will not be discussed in this document. Other researchers have reported on semi-span transonic testing correction issues, however most of the reported research has been for low Mach testing. Some of the reports for low Mach testing address the difficult problem of preventing undesirable airflow under a semi-span model while ensuring unrestricted main balance functionality, however, for transonic models this issue has gone unresolved. BCA has been cognizant for sometime that there are marked differences in wing pressure distributions from semi-span transonic model testing than from full model or flight testing. It has been suspected that these differences are at least in part due to airflow under the model. Previous efforts by BCA to address this issue have proven to be ineffective or inconclusive and in one situation resulted in broken hardware. This paper reports on a Boeing-NASA collaborative investigation based on a series of small tests conducted between June 2006 and November 2007 in the 11 by 11 foot Transonic Wind Tunnel at NASA Ames on three large commercial jet
Kennelly, Robert A., Jr.; Kroo, Ilan M.; Strong, James M.; Carmichael, Ralph L.
An experimental investigation was conducted at the ARC 11- by 11-Foot Transonic Wind Tunnel as part of the Oblique Wing Research Aircraft Program to study the aerodynamic performance and stability characteristics of a 0.087-scale model of an F-8 airplane fitted with an oblique wing designed by Rockwell International. The 10.3 aspect ratio, straight-tapered wing of 0.14 thickness/chord ratio was tested at two different mounting heights above the fuselage. Additional tests were conducted to assess low-speed behavior with and without flaps, aileron effectiveness at representative flight conditions, and transonic drag divergence with 0 degree wing sweep. Longitudinal stability data were obtained at sweep angles of 0, 30, 45, 60, and 65 degrees, at Mach numbers ranging from 0.25 to 1.40. Test Reynolds numbers varied from 3.2 to 6.6 x 10 exp 6/ft. and angle of attack ranged from -5 to +18 degrees. Most data were taken at zero sideslip, but a few runs were at sideslip angles of +/- 5 degrees. The raised wing position proved detrimental overall, although side force and yawing moment were reduced at some conditions. Maximum lift coefficient with the flaps deflected was found to fall short of the value predicted in the preliminary design document. The performance and trim characteristics of the present wing are generally inferior to those obtained for a previously tested wing designed at ARC.
Various papers on unsteady transonic aerodynamics are presented. The topics addressed include: physical phenomena associated with unsteady transonic flows, basic equations for unsteady transonic flow, practical problems concerning aircraft, basic numerical methods, computational methods for unsteady transonic flows, application of transonic flow analysis to helicopter rotor problems, unsteady aerodynamics for turbomachinery aeroelastic applications, alternative methods for modeling unsteady transonic flows.
Bodapati, S.; Lee, C.-S.
The unsteady wake profiles of an airfoil with an oscillating flap were measured in the NASA Ames 11 x 11-foot transonic wind tunnel. Laser Doppler Velocimetry (LDV) and holography techniques were used in limited region where optical accessability is available. X-hot-film wire was used to measure the wake profiles in the complete region to obtain magnitude and direction of the flow. A thorough calibration was carried out to determine the sensitivity coefficients of the hot-wire in three different tunnels at transonic speeds. A calculation procedure is established to resolve the hot-wire signals at transonic speeds and applied in the measurements of steady and periodic wake profiles. The effect of flow incidence on the hot-wire signals is evaluated and incorporated in the analyses. Typical hot-wire results are compared with the results of LDV, holography and pitot-static tube embedded with Kulite transducers.
Benser, W. A.
The highlights of the NASA program on transonic compressors are presented. Effects of blade shape and throat area on losses and flow range are discussed. Some effects of casing treatment on stall margin are presented. Results of tests with varying solidity are also presented. High Mach number, highly loaded stators are discussed and some results of stator hub slit suction are presented.
Garabedian, P. R.
Computer codes for the design and analysis of transonic airfoils are considered. The design code relies on the method of complex characteristics in the hodograph plane to construct shockless airfoil. The analysis code uses artificial viscosity to calculate flows with weak shock waves at off-design conditions. Comparisons with experiments show that an excellent simulation of two dimensional wind tunnel tests is obtained. The codes have been widely adopted by the aircraft industry as a tool for the development of supercritical wing technology.
Solomon, George E
Experimental results are presented for transonic flow post cone-cylinder, axially symmetric bodies. The drag coefficient and surface Mach number are studied as the free-stream Mach number is varied and, wherever possible, the experimental results are compared with theoretical predictions. Interferometric results for several typical flow configurations are shown and an example of shock-free supersonic-to-subsonic compression is experimentally demonstrated. The theoretical problem of transonic flow past finite cones is discussed briefly and an approximate solution of the axially symmetric transonic equations, valid for a semi-infinite cone, is presented.
Baals, D. D.
The inability of existing wind tunnels to provide aerodynamic test data at transonic speeds and flight Reynolds numbers was examined. The proposed transonic facility is a high Reynolds number transonic wind tunnel designed to meet the research and development needs of industry, and the scientific community. The facility employs the cryogenic approach to achieve high transonic Reynolds numbers at acceptable model loads and tunnel power. By using temperature as a test variable, a unique capability to separate scale effects from model aeroelastic effects is provided. The performance envelope of the facility is shown to provide a ten fold increase in transonic Reynolds number capability compared to currently available facilities.
Habel, Louis W; Henderson, James H; Miller, Mason F
Report describes the development of the Langley annular transonic tunnel, a facility in which test Mach numbers from 0.6 to slightly over 1.0 are achieved by rotating the test model in an annular passage between two concentric cylinders. Data obtained for two-dimensional airfoil models in the Langley annular transonic tunnel at subsonic and sonic speeds are shown to be in reasonable agreement with experimental data from other sources and with theory when comparisons are made for nonlifting conditions or for equal normal-force coefficients rather than for equal angles of attack. The trends of pressure distributions obtained from measurements in the Langley annular transonic tunnel are consistent with distributions calculated for Prandtl-Meyer flow.
Joh, C.-Y.; Grossman, B.; Haftka, R. T.
Numerical optimization procedures were considered for the design of airfoils in transonic flow based on the transonic small disturbance (TSD) and Euler equations. A sequential approximation optimization technique was implemented with an accurate approximation of the wave drag based on the Nixon's coordinate straining approach. A modification of the Euler surface boundary conditions was implemented in order to efficiently compute design sensitivities without remeshing the grid. Two effective design procedures producing converged designs in approximately 10 global iterations were developed: interchanging the role of the objective function and constraint and the direct lift maximization with move limits which were fixed absolute values of the design variables.
Amaya, Max A.
New baseline turbulence levels have been measured using a new CTA and new hot-wire sensors. Levels remain the same as measured in 1999. Data and methodology documented (almost). New baseline acoustics levels have been measured up to Mach 1.35. -Levels are higher than reported in 1999. -Data and methodology documented (almost). Application of fairings to the strut trailing edge showed up to a 10% reduction in the tunnel background noise. Data analysis and documentation for publishing is ongoing.
Spreiter, J. R.; Stahara, S. S.
The paper presents a review of the historical development in unsteady transonic aerodynamics, along with the foundations and accomplishments of several approaches to solve the equations of unsteady transonic flow. The discussion covers the linearized unsteady flow theory, numerical solution of the exact equations for an inviscid compressible gas, nonlinear small disturbance theory of transonic flow and linearization of the unsteady component about the nonlinear solution for the steady state, local linearization solution for unsteady transonic flow, unsteady transonic flow theory for slender wings and bodies, and three-dimensional unsteady transonic flows. The relation between the calculated results and experiment is examined. It is shown that the newly emerging numerical methods are capable of solving the nonlinear equations for two-dimensional flow and can be extended to three-dimensional flows.
Swisshelm, J. M.; Adamczyk, J. J.
A semidirect method, driven by a Poisson solver, was developed for inviscid transonic flow computations. It is an extension of a recently introduced algorithm for solving subsonic rotational flows. Shocks are captured by implementing a form of artificial compressibility. Nonisentropic cases are computed using a shock tracking procedure coupled with the Rankine-Hugoniot relationships. Results are presented for both subsonic and transonic flows. For the test geometry, an unstaggered cascade of 20 percent thick circular arc airfoils at zero angle of attack, shocks are crisply resolved in supercritical situations and the algorithm converges rapidly. In addition, the convergence rate appears to be nearly independent of the entropy and vorticity production at the shock.
Tinoco, E. N.
The use of computational methods for three dimensional transonic flow design and analysis at the Boeing Company is presented. A range of computational tools consisting of production tools for every day use by project engineers, expert user tools for special applications by computational researchers, and an emerging tool which may see considerable use in the near future are described. These methods include full potential and Euler solvers, some coupled to three dimensional boundary layer analysis methods, for transonic flow analysis about nacelle, wing-body, wing-body-strut-nacelle, and complete aircraft configurations. As the examples presented show, such a toolbox of codes is necessary for the variety of applications typical of an industrial environment. Such a toolbox of codes makes possible aerodynamic advances not previously achievable in a timely manner, if at all.
Surampudi, S. P.; Adamczyk, J. J.
There is a need for methods to predict the unsteady air loads associated with flutter of turbomachinery blading at transonic speeds. The results of such an analysis in which the steady relative flow approaching a cascade of thin airfoils is assumed to be transonic, irrotational, and isentropic is presented. The blades in the cascade are allowed to undergo a small amplitude harmonic oscillation which generates a small unsteady flow superimposed on the existing steady flow. The blades are assumed to oscillate with a prescribed motion of constant amplitude and interblade phase angle. The equations of motion are obtained by linearizing about a uniform flow the inviscid nonheat conducting continuity and momentum equations. The resulting equations are solved by employing the Weiner Hopf technique. The solution yields the unsteady aerodynamic forces acting on the cascade at Mach number equal to 1. Making use of an unsteady transonic similarity law, these results are compared with the results obtained from linear unsteady subsonic and supersonic cascade theories. A parametric study is conducted to find the effects of reduced frequency, solidity, stagger angle, and position of pitching axis on the flutter.
Rubesin, Morris W.; Viegas, John R.
Predictions of several models compared with measurements of well-documented flow. Report reviews performances of variety of mathematical models of turbulence in transonic flow. Predictions of models compared with measurements of flow in wind tunnel along outside of cylinder having axisymmetric bump of circular-arc cross section in meridional plane. Review part of continuing effort to calibrate and verify computer codes for prediction of transonic flows about airfoils. Johnson-and-King model proved superior in predicting transonic flow over bumpy cylinder.
Garabedian, P. R.
The paper develops a method for calculating wave drag for two-dimensional transonic flow, with particular application to the prediction of the drag rise Mach number of a supercritical wing section. The method is based on a transonic similarity model which is defined by a normalized small perturbation equation and represents shock waves by the addition of an artificial viscosity term in the region of supersonic flow to the partial differential equation. The drag formula obtained allows the computer simulation of transonic wind tunnel data taking account of boundary layer and wall effects.
Purcell, Timothy W.
A computational method which predicts far-field impulsive noise from a transonic motor blade is demonstrated. This method couples near-field results from a full-potential finite-difference method with a new Kirchhoff integral formulation to extend the finite-difference results into the acoustic far-field. This Kirchhoff formula is written in a blade-fixed coordinate system. It requires initial data from the potential code on a plane at the sonic radius. A recent hovering rotor experiment is described where accurate pressure measurements were recorded on the sonic cylinder and at 2 and 3 radii. The potential code prediction of sonic cylinder pressures is excellent. Acoustic far-field pressure predictions show good agreement with hover experimental data over the range of speeds from 0.85 to 0.92 tip Mach number, the latter of which have delocalized transonic flow. These results are some of the first successful predictions for peak pressure amplitudes using a computational code.
Horstman, C. C.; Rose, W. C.
The use of hot-wire anemometry for obtaining fluctuating data in transonic flows has been evaluated. From hot-wire heat loss correlations based on previous transonic data, the sensitivity coefficients for velocity, density, and total temperature fluctuations have been calculated for a wide range of test conditions and sensor parameters. For sensor Reynolds numbers greater than 20 and high sensor overheat ratios, the velocity sensitivity remains independent of Mach number and equal to the density sensitivity. These conclusions were verified by comparisons of predicted sensitivities with those from recent direct calibrations in transonic flows. Based on these results, techniques are presented to obtain meaningful measurements of fluctuating velocity, density, and Reynolds shear stress using hot-wire and hot-film anemometers. Examples of these measurements are presented for two transonic boundary layers.
Horstman, C. C.; Rose, W. C.
The use of hot-wire anemometry for obtaining fluctuating data in transonic flows has been evaluated. From hot-wire heat loss correlations based on previous transonic data, the sensitivity coefficients for velocity, density, and total temperature fluctuations have been calculated for a wide range of test conditions and sensor parameters. For sensor Reynolds number greater than 20 and high sensor overheat ratios, the velocity sensitivity remains independent of Mach number and equal to the density sensitivity. These conditions were verified by comparisons of predicted sensitivities with those from recent direct calibrations in transonic flows. Based on these results, techniques are presented to obtain meaningful measurements of fluctuating velocity, density, and Reynolds shear stress using hot-wire and hot-film anemometers. Example of these measurements are presented for two transonic boundary layers.
Guazzotto, Luca; Hameiri, Eliezer
A linear, two-dimensional model of a transonic plasma flow in equilibrium is constructed and given an explicit solution in the form of a complex Laplace integral. The solution indicates that the transonic state can be solved as an elliptic boundary value problem, as is done in the numerical code FLOW [Guazzotto et al., Phys. Plasmas 11, 604 (2004)]. Moreover, the presence of a hyperbolic region does not necessarily imply the presence of a discontinuity or any other singularity of the solution.
Carlson, L. A.
A numerical technique for designing transonic airfoils having a prescribed pressure distribution (the inverse problem) is presented. The method employs the basic features of Jameson's iterative solution for the full potential equation, except that inverse boundary conditions and Cartesian coordinates are used. The method is a direct-inverse approach that controls trailing-edge closure. Examples show the application of the method to design aft-cambered and other airfoils specifically for transonic flight.
Liang, E. P. T.; Thompson, K. A.
The solution for the radial drift velocity of thin disk accretion onto black holes must be transonic, and is analogous to the critical solution in spherical Bondi accretion, except for the presence of angular momentum. The transonic requirement yields a correct treatment of the inner region of the disk not found in the conventional Keplerian models and may lead to significantly different overall disk structures. Possible observational consequences, relevant to the black hole hypothesis for Cyg X-1 and other candidates, are discussed.
Rosen, B. S.
A computational method developed to provide a transonic analysis for upper/lower surface wing-tip mounted winglets is described. Winglets with arbitrary planform, cant and toe angle, and airfoil section can be modeled. The embedded grid approach provides high flow field resolution and the required geometric flexibility. In particular, coupled Cartesian/cylindrical grid systems are used to model the complex geometry presented by canted upper/lower surface winglets. A new rotated difference scheme is introduced in order to maintain the stability of the small-disturbance formulation in the presence of large spanwise velocities. Wing and winglet viscous effects are modeled using a two-dimensional 'strip' boundary layer analysis. Correlations with wind tunnel and flight test data for three transport configurations are included.
Lewy, S.; Gounet, H.
Models for the noise levels from propellers are discussed, with results compared to in-flight measurements. Methods originally applied to noise from light aircraft are modified and extended to high speed passenger aircraft. Noise emitted from propellers has three components: a monopolar emission due to the air displaced by a blade; a bipolar form from average and fluctuating forces exerted by the blades; and a quadripolar component produced by deformation of the streamlines around the blade profile and defined by the Lighthill tensor. The latter is not a factor in the subsonic regime and can be neglected. Attention is given to a formalism which accounts for the sound level along each band, the frequency harmonics at each blade passage, the number of blades, and the rotation rate. The measured directivities of the two components are described. It is found that the radiated noise levels can be reduced in slow aircraft by lowering the peripheral velocity while keeping the same power with more blades. Calculations including the quadripolar term are necessary for modeling noise levels in transonic propellers.
Rubesin, Morris W.; Viegas, John R.
A review is made of the performance of a variety of turbulence models in the evaluation of a particular well documented transonic flow. This is done to supplement a previous attempt to calibrate and verify transonic airfoil codes by including many more turbulence models than used in the earlier work and applying the calculations to an experiment that did not suffer from uncertainties in angle of attack and was free of wind tunnel interference. It is found from this work, as well as in the earlier study, that the Johnson-King turbulence model is superior for transonic flows over simple aerodynamic surfaces, including moderate separation. It is also shown that some field equation models with wall function boundary conditions can be competitive with it.
Chow, Chuen-Yen; Chang, I-Chung; Gea, Lie-Mine
A numerical method is presented for calculating the unsteady transonic rotor flow with aeroelasticity effects. The blade structural dynamic equations based on beam theory were formulated by FEM and were solved in the time domain, instead of the frequency domain. For different combinations of precone, droop, and pitch, the correlations are very good in the first three flapping modes and the first twisting mode. However, the predicted frequencies are too high for the first lagging mode at high rotational speeds. This new structure code has been coupled into a transonic rotor flow code, TFAR2, to demonstrate the capability of treating elastic blades in transonic rotor flow calculations. The flow fields for a model-scale rotor in both hover and forward flight are calculated. Results show that the blade elasticity significantly affects the flow characteristics in forward flight.
Edwards, J. W.; Bland, S. R.; Seidel, D. A.
Comparisons of calculated and experimental transonic unsteady pressures and airloads for four of the AGARD Two Dimensional Aeroelastic Configurations and for a rectangular supercritical wing are presented. The two dimensional computer code, XTRAN2L, implementing the transonic small perturbation equation was used to obtain results for: (1) pitching oscillations of the NACA 64A010A; NLR 7301 and NACA 0012 airfoils; (2) flap oscillations for the NACA 64A006 and NRL 7301 airfoils; and (3) transient ramping motions for the NACA 0012 airfoils. Results from the three dimensional code XTRAN3S are compared with data from a rectangular supercritical wing oscillating in pitch. These cases illustrate the conditions under which the transonic inviscid small perturbation equation provides reasonable predictions.
Carlson, L. A.
A numerical technique was developed for the analysis of specified transonic airfoils or for the design of airfoils having a prescribed pressure distribution, including the effect of weak viscous interaction. The method uses the full potential equation, a stretched Cartesian coordinate system, and the Nash-MacDonald turbulent boundary layer method. Comparisons with experimental data for typical transonic airfoils show excellent agreement. An example shows the application of the method to design a thick aft-cambered airfoil, and the effects of viscous interaction on its performance are discussed.
Bruce, Robert A.; Hess, Robert W.; Rivera, Jose A., Jr.
A vapor generator was developed for use in the NASA Langley Transonic Dynamics Tunnel (TDT). Propylene glycol was used as the vapor material. The vapor generator system was evaluated in a laboratory setting and then used in the TDT as part of a laser light sheet flow visualization system. The vapor generator provided satisfactory seeding of the air flow with visible condensate particles, smoke, for tests ranging from low subsonic through transonic speeds for tunnel total pressures from atmospheric pressure down to less than 0.1 atmospheric pressure.
Abramowicz, Marek A.; Kato, Shoji
Regularity conditions and global topological constraints leave some forbidden regions in the parameter space of the transonic isothermal, rotating matter onto black holes. Unstable flows occupy regions touching the boundaries of the forbidden regions. The astrophysical consequences of these results are discussed.
Reed, T. D.; Pope, T. C.; Cooksey, J. M.
State-of-the art instrumentation and procedures for calibrating transonic (0.6 less than M less than 1.4) and supersonic (M less than or equal to 3.5) wind tunnels were reviewed and evaluated. Major emphasis was given to transonic tunnels. Continuous, blowdown and intermittent tunnels were considered. The required measurements of pressure, temperature, flow angularity, noise and humidity were discussed, and the effects of measurement uncertainties were summarized. A comprehensive review of instrumentation currently used to calibrate empty tunnel flow conditions was included. The recent results of relevant research are noted and recommendations for achieving improved data accuracy are made where appropriate. It is concluded, for general testing purposes, that satisfactory calibration measurements can be achieved in both transonic and supersonic tunnels. The goal of calibrating transonic tunnels to within 0.001 in centerline Mach number appears to be feasible with existing instrumentation, provided correct calibration procedures are carefully followed. A comparable accuracy can be achieved off-centerline with carefully designed, conventional probes, except near Mach 1. In the range 0.95 less than M less than 1.05, the laser Doppler velocimeter appears to offer the most promise for improved calibration accuracy off-centerline.
Bobbitt, C., Jr.; Everhart, J.
This paper describes the current activities at the National Transonic Facility to document the test-section flow and to support tunnel improvements. The paper is divided into sections on the tunnel calibration, flow quality measurements, data quality assurance, and implementation of wall interference corrections.
Davis, Sanford S.; Malcolm, Gerald N.
Unsteady aerodynamic loads were measured on an oscillating NACA 64A010 airfoil In the NASA Ames 11 by 11 ft Transonic Wind Tunnel. Data are presented to show the effect of the unsteady shock-wave/boundary-layer interaction on the fundamental frequency lift, moment, and pressure distributions. The data show that weak shock waves induce an unsteady pressure distribution that can be predicted quite well, while stronger shock waves cause complex frequency-dependent distributions due to flow separation. An experimental test of the principles of linearity and superposition showed that they hold for weak shock waves while flows with stronger shock waves cannot be superimposed.
Harder, Keith C; Klunker, E B
The basic ideas of the slender-body approximation have been applied to the nonlinear transonic-flow equation for the velocity potential in order to obtain some of the essential features of slender-body theory at transonic speeds. The results of the investigation are presented from a unified point of view which demonstrates the similarity of slender-body solutions in the various Mach number ranges. The transonic area rule and some conditions concerning its validity follow from the analysis. (author)
Foughner, Jerome T., Jr. (Compiler)
In order to assess the state of the art in transonic flow disciplines and to glimpse at future directions, NASA-Langley held a Transonic Symposium. Emphasis was placed on steady, three dimensional external, transonic flow and its simulation, both numerically and experimentally. The symposium included technical sessions on wind tunnel and flight experiments; computational fluid dynamic applications; inviscid methods and grid generation; viscous methods and boundary layer stability; and wind tunnel techniques and wall interference. This, being volume 1, is unclassified.
Foughner, Jerome T., Jr. (Compiler)
Papers presented at the Transonic Symposium are compiled. The following subject areas are covered: National Transonic Facility status; transonic aerodynamics of slender wing-body configuration; laminar flow flight experiments; laminar flow wind tunnel experiments; computational support of X-29A flight experiment; transition location on a clean-up glove installed on a F-14 aircraft; and design studies for a laminar glove for the X-29 aircraft.
Bechert, D. W.; Hage, W.; Stanewsky, E.
Wave drag reduction bodies on the suction side of transonic wings are investigated. Following the original invention by O. Frenzl (1942), subsequently, such bodies have been suggested by Kuechemann and Whitcomb. These devices have been used sucessfully on various TUPOLEV aircraft and on the CONVAIR 990 airliner. New transonic wind tunnel data from an unswept wing with an array of Kuechemann Carrots are presented (airfoil: CAST 10/DOA-2). In a certain parameter range (M= 0.765-0.86) the measurements exhibit a significant reduction of the shock strength on a wing between the Kuechemann Carrots. This entails a dramatic reduction of drag, in a certain Mach number and angular regime up to 50-60%.
A review of recent and historical work in the field of transonic and supersonic ground effect aerodynamics has been conducted, focussing on applied research on wings and aircraft, present and future ground transportation, projectiles, rocket sleds and other related bodies which travel in close ground proximity in the compressible regime. Methods for ground testing are described and evaluated, noting that wind tunnel testing is best performed with a symmetry model in the absence of a moving ground; sled or rail testing is ultimately preferable, though considerably more expensive. Findings are reported on shock-related ground influence on aerodynamic forces and moments in and accelerating through the transonic regime - where force reversals and the early onset of local supersonic flow is prevalent - as well as more predictable behaviours in fully supersonic to hypersonic ground effect flows.
Edwards, John W.
Since coming online in 1984, the National Transonic Facility (NTF) cryogenic wind tunnel at the NASA Langley Research Center has provided unique high Reynolds number testing capability. While turbulence levels in the tunnel, expressed in terms of percent dynamic pressure, are typical of other transonic wind tunnels, the significantly increased load levels utilized to achieve flight Reynolds numbers, in conjunction with the unique structural design requirements for cryogenic operation, have brought forward the issue of model and model support structure vibrations. This paper reports new experimental measurements documenting aerodynamic and structural dynamics processes involved in such vibrations experienced in the NTF. In particular, evidence of local un-steady airloads developed about the model support strut is shown and related to well documented acoustic features known as "Parker" modes. Two-dimensional unsteady viscous computations illustrate this model support structure loading mechanism.
Young, Clarence P., Jr.; Hergert, Dennis W.; Butler, Thomas W.; Herring, Fred M.
A buffet test of a commercial transport model was accomplished in the National Transonic Facility at the NASA Langley Research Center. This aeroelastic test was unprecedented for this wind tunnel and posed a high risk to the facility. This paper presents the test results from a structural dynamics and aeroelastic response point of view and describes the activities required for the safety analysis and risk assessment. The test was conducted in the same manner as a flutter test and employed onboard dynamic instrumentation, real time dynamic data monitoring, automatic, and manual tunnel interlock systems for protecting the model. The procedures and test techniques employed for this test are expected to serve as the basis for future aeroelastic testing in the National Transonic Facility. This test program was a cooperative effort between the Boeing Commercial Airplane Company and the NASA Langley Research Center.
Lores, M. E.; Burdges, K. P.; Shrewsbury, G. D.
Numerical optimization was used in conjunction with an inviscid, full potential equation, transonic flow analysis computer code to design an upper surface contour for a conventional airfoil to improve its supercritical performance. The modified airfoil was tested in a compressible flow wind tunnel. The modified airfoil's performance was evaluated by comparison with test data for the baseline airfoil and for an airfoil developed by optimization of leading edge of the baseline airfoil. While the leading edge modification performed as expected, the upper surface re-design did not produce all of the expected performance improvements. Theoretical solutions computed using a full potential, transonic airfoil code corrected for viscosity were compared to experimental data for the baseline airfoil and the upper surface modification. These correlations showed that the theory predicted the aerodynamics of the baseline airfoil fairly well, but failed to accurately compute drag characteristics for the upper surface modification.
Verhoff, Vincent G.; Camperchioli, William P.; Lopez, Isaac
NASA LeRC has designed and constructed a new state-of-the-art test facility. This facility, the Transonic Turbine Blade Cascade, is used to evaluate the aerodynamics and heat transfer characteristics of blade geometries for future turbine applications. The facility's capabilities make it unique: no other facility of its kind can combine the high degree of airflow turning, infinitely adjustable incidence angle, and high transonic flow rates. The facility air supply and exhaust pressures are controllable to 16.5 psia and 2 psia, respectively. The inlet air temperatures are at ambient conditions. The facility is equipped with a programmable logic controller with a capacity of 128 input/output channels. The data acquisition system is capable of scanning up to 1750 channels per sec. This paper discusses in detail the capabilities of the facility, overall facility design, instrumentation used in the facility, and the data acquisition system. Actual research data is not discussed.
Kilgore, W. A.; Balakrishna, S.; Bobbitt, C. W.; Underwood, P.
The National Transonic Facility continues to make enhancements to provide quality data in a safe, efficient and cost effective method for aerodynamic ground testing. Recent enhancements discussed in this paper include the restoration of reliability and improved performance of the heat exchanger systems resulting in the expansion of the NTF air operations envelope. Additionally, results are presented from a continued effort to reduce model dynamics through the use of a new stiffer balance and sting.
Kilgore, W. A.; Balakrishna, S.; Bobbitt, C. W.; Underwood, P.
The National Transonic Facility continues to make enhancements to provide quality data in a safe, efficient and cost effective method for aerodynamic ground testing. Recent enhancements discussed in this paper include the restoration of reliability and improved performance of the heat exchanger systems resulting in the expansion of the NTF air operations envelope. Additionally, results are presented from a continued effort to reduce model dynamics through the use of a new stiffer balance and sting
the difficulties related to transonic testing techniques. The presence of a model mounting system, proximity of the tunnel walls, the method and amount...avec la methode exposee rtf 2. (b) Le coefficient de poussee interne KTJ , qui represente le precedent exprime en pressions relatives PQAJ KTi... method was developed to define a valid total pressure, based on a mast flow averaging procedure, for a distorted jet pipe flow. The results for the AGARD
Nagai, Naonori; Iwatani, Junji
Sectional profiles close to a tip 124 and a part between a midportion 125 and a hub 123 are shifted to the upstream of an operating fluid flow in a sweep direction. Accordingly, an S shape is formed in which the tip 124 and the part between the midportion 125 and the hub 123 protrude. As a result, it is possible reduce various losses due to shook, waves, thereby forming a transonic airfoil having an excellent aerodynamic characteristic.
RECEIVED JUL 0 12005 FINAL REPORT FOR: AFOSR GRANT F49260-02-1-0284 TRANSONIC CASCADE MEASUREMENTS TO SUPPORT ANALYTICAL MODELING Paul A. Durbin ...PAD); 650-723-1971 (JKE) durbin @vk.stanford.edu; email@example.com submitted to: Attn: Dr. John Schmisseur Air Force Office of Scientific Research...both spline and control points for subsequent wall shape definitions. An algebraic grid generator was used to generate the grid for the blade-wall
Seginer, A.; Rosenwasser, I.
Magnus forces and moments were measured on a basic-finner model spinning in transonic flow. Spin was induced by canted fins or by full-span or semi-span, outboard and inboard roll controls. Magnus force and moment reversals were caused by Mach number, reduced spin rate, and angle of attack variations. Magnus center of pressure was found to be independent of the angle of attack but varied with the Mach number and model configuration or reduced spin rate.
Jameson, Antony; Ou, Kui
This article traces the evolution of long range jet transport aircraft over the 50 years since Kuechemann founded the journal Progress in Aerospace Sciences. The article is particularly focused on transonic aerodynamics. During Kuechemann's life time a good qualitative understanding had been achieved of transonic flow and swept wing design, but transonic flow remained intractable to quantitative prediction. During the last 50 years this situation has been completely transformed by the introduction of sophisticated numerical algorithms and an astonishing increase in the available computational power, with the consequence that aerodynamic design is now carried out largely by computer simulation. Moreover developments in aerodynamic shape optimization based on control theory enable a competitive swept wing to be designed in just two simulations, as illustrated in the article. While the external appearance of long range jet aircraft has not changed much, advances in information technology have actually transformed the entire design and manufacturing process through parallel advances in computer aided design (CAD), computational structural mechanics (CSM) and multidisciplinary optimization (MDO). They have also transformed aircraft operations through the adoption of digital fly-by-wire and advanced navigational techniques.
Baurdoux, H.I. Symmetrical Transonic Potential Flows Around Quasi-Elliptical Airfoil Boerstoel , J.W. Sections. NLR-TR69007U, National Aerospace...Laboratory, The Netherlands, December 1968. 14. Kacprzynski, J.J. Wind Tunnel Tests of a Boerstoel Shockless Symmetrical Airfoil. NRC, NAE, 5x5 ft Transonic
Berry, H. M.; Batina, J. T.; Yang, T. Y.
Viscous effects on transonic airfoil stability and response are investigated using an integral boundary layer model coupled to the inviscid XTRAN2L transonic small disturbance code. Unsteady transonic airloads required for stability analyses are computed using a pulse transfer function analysis including viscous effects. The pulse analysis provides unsteady aerodynamic forces for a wide range of reduced frequency in a single flow field computation. Nonlinear time marching aeroelastic solutions are presented which show the effects of viscosity on airfoil response behavior and flutter. Effects of amplitude on time marching responses are demonstrated. A state space aeroelastic model employing Pade approximants to describe the unsteady airloads is used to study the effects of viscosity on transonic airfoil stability. State space dynamic pressure root loci are in good overall agreement with time marching damping and frequency estimates. Parallel sets of results with and without viscous effects reveal the effects of viscosity on transonic unsteady airloads and aeroelastic characteristics of airfoils.
Tumkur Karnick, Pradeepa; Venkatraman, Kartik
The transonic flutter dip of an aeroelastic system is primarily caused by compressibility of the flowing fluid. Viscous effects are not dominant in the pre-transonic dip region. In fact, an Euler solver can predict this flutter boundary with considerable accuracy. However with an increase in Mach number the shock moves towards the trailing edge causing shock induced separation. This shock-boundary layer interaction changes the flutter boundary in the transonic and post-transonic dip region significantly. We discuss the effect of viscosity in changing the flutter boundary in the post-transonic dip region using a RANS solver coupled to a two-degree of freedom model of the structural dynamics of a wing.
Young, Clarence P., Jr.; Hergert, Dennis W.; Butler, Thomas W.; Herring, Fred M.
A buffet test of a commercial transport model was accomplished in the National Transonic Facility at the NASA Langley Research Center. This aeroelastic test was unprecedented for this wind tunnel and posed a high risk for the facility. Presented here are the test results from a structural dynamics and aeroelastic response point of view. The activities required for the safety analysis and risk assessment are described. The test was conducted in the same manner as a flutter test and employed on-board dynamic instrumentation, real time dynamic data monitoring, and automatic and manual tunnel interlock systems for protecting the model.
Carlson, L. A.
A numerical technique for analyzing transonic airfoils is presented. The method employs the basic features of Jameson's iterative solution for the full potential equation, except that Cartesian coordinates are used rather than a grid which fits the airfoil, such as the conformal circle-plane or 'sheared parabolic' coordinates which were used previously. Comparison with previous results shows that it is not necessary to match the computational grid to the airfoil surface, and that accurate results can be obtained with a Cartesian grid for lifting supercritical airfoils.
Bendiksen, Oddvar O.; Kousen, Kenneth A.
In transonic flutter problems where shock motion plays an important part, it is believed that accurate predictions of the flutter boundaries will require the use of codes based on the Euler equations. Only Euler codes can obtain the correct shock location and shock strength, and the crucially important shock excursion amplitude and phase lag. The present study is based on the finite volume scheme developed by Jameson and Venkatakrishnan for the 2-D unsteady Euler equations. The equations are solved in integral form on a moving grid. The variable are pressure, density, Cartesian velocity components, and total energy.
hodograph method of Boerstoel (defs. 1, 2). While the airfoil was oscillating in pitch about an axis at 40 per cent of the chord detailed pressure...Aerospace Programs (NIVR). The authors are indebted to the NIVR and Fokker-VFW for the permission to use this material. 7. REFERENCES 1 Boerstoel , J.W...and Transonic shock-free aerofoil design by an analytic Huizing, G.H. hodograph method Journal of Aircraft, 12 No. 9, pp 730-736, 1975, 2 Boerstoel
wing chord is 117 mm, the span is 47.5 mm, and the inlet and exit heights are 39.7 and 18.7 mm respectively. The rest of the wind tunnel was designed...around the geometry to adapt it to our steady-flow, transonic wind tunnel used previously in this program, while maximizing optical access to the wing ...the wing surface isentropic Mach number distribution using equation 4.1. 4.4 Overall Wind Tunnel System Flow is provided by an Ingersoll-Rand
Lee, J. D.; Gregorek, G. M.
A limited-zone ventilated wall panel was developed for a closed-wall icing tunnel which permitted correct simulation of transonic flow over model rotor airfoil sections with and without ice accretions. Candidate porous panels were tested in the Ohio State University 6- x 12-inch transonic airfoil tunnel and result in essentially interference-free flow, as evidenced by pressure distributions over a NACA 0012 airfoil for Mach numbers up to 0.75. Application to the NRC 12- x 12-inch icing tunnel showed a similar result, which allowed proper transonic flow simulation in that tunnel over its full speed range.
Beam, R. M.; Warming, R. F.
The feasibility of obtaining two-dimensional, unsteady transonic aerodynamic data by numerically integrating the Euler equations is investigated. An explicit, third-order-accurate, noncentered, finite-difference scheme is used to compute unsteady flows about airfoils. Solutions for lifting and nonlifting airfoils are presented and compared with subsonic linear theory. The applicability and efficiency of the numerical indicial function method are outlined. Numerically computed subsonic and transonic oscillatory aerodynamic coefficients are presented and compared with those obtained from subsonic linear theory and transonic wind-tunnel data.
Adamson, T. C., Jr.; Messiter, A. F.
For many internal transonic flows of practical interest, some of the relevant nondimensional parameters typically are small enough that a perturbation scheme can be expected to give a useful level of numerical accuracy. A variety of steady and unsteady transonic channel and cascade flows is studied with the help of systematic perturbation methods which take advantage of this fact. Asymptotic representations are constructed for small changes in channel cross-section area, small flow deflection angles, small differences between the flow velocity and the sound speed, small amplitudes of imposed oscillations, and small reduced frequencies. Inside a channel the flow is nearly one-dimensional except in thin regions immediately downstream of a shock wave, at the channel entrance and exit, and near the channel throat. A study of two-dimensional cascade flow is extended to include a description of three-dimensional compressor-rotor flow which leads to analytical results except in thin edge regions which require numerical solution. For unsteady flow the qualitative nature of the shock-wave motion in a channel depends strongly on the orders of magnitude of the frequency and amplitude of impressed wall oscillations or fluctuations in back pressure. One example of supersonic flow is considered, for a channel with length large compared to its width, including the effect of separation bubbles and the possibility of self-sustained oscillations. The effect of viscosity on a weak shock wave in a channel is discussed.
Isom, Morris; Purcell, Timothy W.; Strawn, Roger C.
A new method is presented for predicting the impulsive noise generated by a transonic rotor blade. The method is a combined approach involving computational fluid dynamics and geometrical acoustics. A full-potential finite-difference method is used to obtain the pressure field close to the blade. A Kirchhoff integral formulation is then used to extend these finite-difference results into the far field. This Kirchhoff formula is based on geometrical acoustics approximations. It requires initial data across a plane at the sonic radius in a blade-fixed coordinate system. This data is provided by the finite-difference solution. Acoustic pressure predictions show good agreement with hover experimental data for cases with hover tip Mach numbers of 0.88 through 0.96. The cases above 0.92 tip Mach number are dominated by non-linear transonic effects seen as strong shocks on and off the blade tip. This paper gives the first successful predictions of far-field acoustic pressures for high-speed impulsive noise over a range of Mach numbers after delocalization.
Chaparro, Daniel; Fujiwara, Gustavo E. C.; Ting, Eric; Nguyen, Nhan
The need to rapidly scan large design spaces during conceptual design calls for computationally inexpensive tools such as the vortex lattice method (VLM). Although some VLM tools, such as Vorview have been extended to model fully-supersonic flow, VLM solutions are typically limited to inviscid, subcritical flow regimes. Many transport aircraft operate at transonic speeds, which limits the applicability of VLM for such applications. This paper presents a novel approach to correct three-dimensional VLM through coupling of two-dimensional transonic small disturbance (TSD) solutions along the span of an aircraft wing in order to accurately predict transonic aerodynamic loading and wave drag for transport aircraft. The approach is extended to predict flow separation and capture the attenuation of aerodynamic forces due to boundary layer viscosity by coupling the TSD solver with an integral boundary layer (IBL) model. The modeling framework is applied to the NASA General Transport Model (GTM) integrated with a novel control surface known as the Variable Camber Continuous Trailing Edge Flap (VCCTEF).
7. VIEW NORTHWEST OF SUBSONIC WIND TUNNEL BUILDING TO TRANSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD
2. VIEW SOUTH OF TRANSONIC WIND TUNNEL BUILDING AND SUPERSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD
5. VIEW NORTHWEST OF SUBSONIC WIND TUNNEL BUILDING TO TRANSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD
1. VIEW SOUTHWEST OF SUBSONIC WIND TUNNEL BUILDING AND TRANSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD
4. VIEW NORTHWEST OF SUPERSONIC WIND TUNNEL BUILDING TO TRANSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD
3. VIEW SOUTHEAST OF TRANSONIC WIND TUNNEL BUILDING TO SUBSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD
Lan, C. Edward
An inviscid transonic code capable of designing an axisymmetric body in a uniform or nonuniform flow was developed. The design was achieved by direct optimiation by coupling an analysis code with an optimizer. Design examples were provided for axisymmetric bodies with fineness ratios of 8.33 and 5 at different Mach numbers. It was shown that by reducing the nose radius and increasing the afterbody thickness of initial shapes obtained from symmetric NACA four-digit airfoil contours, wave drag could be reduced by 29 percent for a body of fineness ratio 8.33 in a nonuniform transonic flow of M = 0.98 to 0.995. The reduction was 41 percent for a body of fineness ratio 5 in a uniform transonic flow of M = 0.925 and 65 percent for the same body but in a nonuniform transonic flow of M = 0.90 to 0.95.
Bland, Samuel R. (Compiler)
Computational fluid dynamics methods have been widely accepted for transonic aeroelastic analysis. Previously, calculations with the TSD methods were used for 2-D airfoils, but now the TSD methods are applied to the aeroelastic analysis of the complete aircraft. The Symposium papers are grouped into five subject areas, two of which are covered in this part: (1) Transonic Small Disturbance (TSD) theory for complete aircraft configurations; and (2) Full potential and Euler equation methods.
King, L. S.; Johnson, D. A.
Navier-Stokes transonic airfoil calculations based on a recently developed nonequilibrium, turbulence closure model are presented for a supercritical airfoil section at transonic cruise conditions and for a conventional airfoil section at shock-induced stall conditions. Comparisons with experimental data are presented which show that this nonequilibrium closure model performs significantly better than the popular Baldwin-Lomax and Cebeci-Smith equilibrium algebraic models when there is boundary-layer separation that results from the inviscid-viscous interactions.
Holst, Terry L.; Pulliam, Thomas H.; Kwak, Dochan (Technical Monitor)
A method for aerodynamic shape optimization based on a genetic algorithm approach is demonstrated. The algorithm is coupled with a transonic full potential flow solver and is used to optimize the flow about transonic wings including multi-objective solutions that lead to the generation of pareto fronts. The results indicate that the genetic algorithm is easy to implement, flexible in application and extremely reliable.
Dougherty, F. C.
The Transonic Airfoil analysis computer code, TAIR, was developed to employ a fast, fully implicit algorithm to solve the conservative full-potential equation for the steady transonic flow field about an arbitrary airfoil immersed in a subsonic free stream. The full-potential formulation is considered exact under the assumptions of irrotational, isentropic, and inviscid flow. These assumptions are valid for a wide range of practical transonic flows typical of modern aircraft cruise conditions. The primary features of TAIR include: a new fully implicit iteration scheme which is typically many times faster than classical successive line overrelaxation algorithms; a new, reliable artifical density spatial differencing scheme treating the conservative form of the full-potential equation; and a numerical mapping procedure capable of generating curvilinear, body-fitted finite-difference grids about arbitrary airfoil geometries. Three aspects emphasized during the development of the TAIR code were reliability, simplicity, and speed. The reliability of TAIR comes from two sources: the new algorithm employed and the implementation of effective convergence monitoring logic. TAIR achieves ease of use by employing a "default mode" that greatly simplifies code operation, especially by inexperienced users, and many useful options including: several airfoil-geometry input options, flexible user controls over program output, and a multiple solution capability. The speed of the TAIR code is attributed to the new algorithm and the manner in which it has been implemented. Input to the TAIR program consists of airfoil coordinates, aerodynamic and flow-field convergence parameters, and geometric and grid convergence parameters. The airfoil coordinates for many airfoil shapes can be generated in TAIR from just a few input parameters. Most of the other input parameters have default values which allow the user to run an analysis in the default mode by specifing only a few input parameters
Wahls, R. A.
An overview of the National Transonic Facility (NTF) from a research utilization perspective is provided. The facility was born in the 1970s from an internationally recognized need for a high Reynolds number test capability based on previous experiences with preflight predictions of aerodynamic characteristics and an anticipated need in support of research and development for future aerospace vehicle systems. Selection of the cryogenic concept to meet the need, unique capabilities of the facility, and the eventual research utilization of the facility are discussed. The primary purpose of the paper is to expose the range of investigations that have used the NTF since being declared operational in late 1984; limited research results are included, though many more can be found in the references.
Srivastava, R.; Bakhle, M. A.; Keith, T. G., Jr.; Stefko, G. L.
This paper describes the calculation of flutter stability characteristics for a transonic forward swept fan configuration using a viscous aeroelastic analysis program. Unsteady Navier-Stokes equations are solved on a dynamically deforming, body fitted, grid to obtain the aeroelastic characteristics using the energy exchange method. The non-zero inter-blade phase angle is modeled using phase-lagged boundary conditions. Results obtained show good correlation with measurements. It is found that the location of shock and variation of shock strength strongly influenced stability. Also, outboard stations primarily contributed to stability characteristics. Results demonstrate that changes in blade shape impact the calculated aerodynamic damping, indicating importance of using accurate blade operating shape under centrifugal and steady aerodynamic loading for flutter prediction. It was found that the calculated aerodynamic damping was relatively insensitive to variation in natural frequency.
Kern, F. A.; Knight, C. W.; Zasimowich, R. F.
The Mach number system for the Langley Research Center's National Transonic Facility was designed to measure pressures to determine Mach number to within + or - 0.002. Nine calibration laboratory type fused quartz gages, four different range gages for the total pressure measurement, and five different range gages for the static pressure measurement were used to satisfy the accuracy requirement over the 103,000-890,000 Pa total pressure range of the tunnel. The system which has been in operation for over 1 year is controlled by a programmable data process controller to select, through the operation of solenoid valves, the proper range fused quartz gage to maximize the measurement accuracy. The pressure gage's analog outputs are digitized by the process controller and transmitted to the main computer for Mach number computation. An automatic two-point on-line calibration of the nine quartz gages is provided using a high accuracy mercury manometer.
Johnson, F. T.; Bussoletti, J. E.; James, R. M.; Young, D. P.; Woo, A. C.
A method is presented that has the potential for solving transonic flow problems about the same complex aircraft configurations currently being analyzed by subsonic panel methods. This method does not require the generation of surface fitted grids. Instead it uses rectangular grids and subgrids together with embedded surface panels on which boundary conditions are imposed. Both the Euler and full potential equations are considered. The method of least squares is used to reduce the solution of these equations to the solution of a sequence of Poisson problems. The Poisson problems are solved using fast Fourier transforms and panel influence coefficient techniques. The overall method is still in its infancy but some two dimensional results are shown illustrating various key features.
Holst, Terry L.; Kwak, Dochan (Technical Monitor)
This presentation describes the state of transonic flow simulation using nonlinear potential methods for external aerodynamic applications. The presentation begins with a review of the various potential equation forms (with emphasis on the full potential equation) and includes a discussion of pertinent mathematical characteristics and all derivation assumptions. Impact of the derivation assumptions on simulation accuracy, especially with respect to shock wave capture, is discussed. Key characteristics of all numerical algorithm types used for solving nonlinear potential equations, including steady, unsteady, space marching, and design methods, are described. Both spatial discretization and iteration scheme characteristics are examined. Numerical results for various aerodynamic applications are included throughout the presentation to highlight key discussion points. The presentation ends with concluding remarks and recommendations for future work. Overall. nonlinear potential solvers are efficient, highly developed and routinely used in the aerodynamic design environment for cruise conditions. Published by Elsevier Science Ltd. All rights reserved.
Wernet, Mark P.
Particle Imaging Velocimetry (PIV) is a powerful measurement technique which can be used as an alternative or complementary approach to Laser Doppler Velocimetry (LDV) in a wide range of research applications. PIV data are measured simultaneously at multiple points in space, which enables the investigation of the non-stationary spatial structures typically encountered in turbomachinery. Many of the same issues encountered in the application of LDV techniques to rotating machinery apply in the application of PIV. Preliminary results from the successful application of the standard 2-D PIV technique to a transonic axial compressor are presented. The lessons learned from the application of the 2-D PIV technique will serve as the basis for applying 3-component PIV techniques to turbomachinery.
Liou, M.-S.; Coakley, T. J.; Bergmann, M. Y.
Numerical simulations were made of two-dimensional transonic flows in diffusers, including flow separation induced by a shock or adverse pressure gradient. The mass-averaged, time-dependent, compressible Navier-Stokes equations, simplified by the thin-layer approximation, were solved using MacCormack's hybrid method. The eddy-viscosity formulation was described by the Wilcox-Rubesin's two-equation, k-omega model. Detailed comparison of the computed results with measurements showed good agreement in all cases, including one with massive separation induced by a strong shock. The computation correctly predicted the details of a distinct lambda shock pattern, closely duplicating the configuration observed experimentally in spark-schlieren photographs.
A method of characterizing the indicial curve for the response frequencies of flutter in the transonic regime by means of a time scale parameter is illustrated. Time linearization of the unsteady potential equation is assumed adequate and the perturbed potential about a steady flow is defined with appropriate boundary conditions. An indicial motion can then be obtained for a given mode, with the lift and drag coefficients which result from a step change in incidence or control surface deflection available to describe any kind of motion. The lift and drag coefficients are shown to result analytically by addition of the multiplicand involving the time scale parameter. If the coefficients are maximized, then at reduced frequencies only two parameters are necessary to form the indicial response curve.
The application of well known design concepts with the combined use of thick transonic profiles to aircraft wing design was investigated. Optimization in terms of weight and operational costs was emphasized. It is shown that the usual design criteria and concepts are too restricted and do not sufficiently represent the physical processes over the wing. Suggestions are made for improving this situation, and a design example given. Compared with a wing design according to previously used criteria, the new design is found to be superior in the most important functions. It is concluded that an isobar concept adjusted to the planform in conjunction with an 'organically' designed wing will lead to the weight optimum solutions of wing profiles.
A method for computing the three-dimensional transonic flow around the blades of a compressor or of a propeller is given. The method is based on the use of the velocity potential, on the hypothesis that the flow is inviscid, irrotational and isentropic. The equation of the potential is solved in a transformed space such that the surface of the blade is mapped into a plane where the periodicity is implicit. This equation is in a nonconservative form and is solved with the help of a finite difference method using artificial time. A computer code is provided and some sample results are given in order to demonstrate the influence of three-dimensional effects and the blade's rotation.
Bae, Myoungjean; Feldman, Mikhail
We establish existence, uniqueness and stability of transonic shocks for a steady compressible non-isentropic potential flow system in a multidimensional divergent nozzle with an arbitrary smooth cross-section, for a prescribed exit pressure. The proof is based on solving a free boundary problem for a system of partial differential equations consisting of an elliptic equation and a transport equation. In the process, we obtain unique solvability for a class of transport equations with velocity fields of weak regularity (non-Lipschitz), an infinite dimensional weak implicit mapping theorem which does not require continuous Fréchet differentiability, and regularity theory for a class of elliptic partial differential equations with discontinuous oblique boundary conditions.
Sancho, Jorge; Vargas, M.; Gonzalez, Ezequiel; Rodriguez, Manuel
In the framework of EXPERT (European Experimental Re-entry Test-bed) accurate transonic aerodynamic coefficients are of paramount importance for the correct trajectory assessment and parachute deployment. A combined CFD (Computational Fluid Dynamics) modelling and experimental campaign strategy was selected to obtain accurate coefficients. A preliminary set of coefficients were obtained by CFD Euler inviscid computation. Then experimental campaign was performed at DNW facilities at NLR. A profound review of the CFD modelling was done lighten up by WTT results, aimed to obtain reliable values of the coefficients in the future (specially the pitching moment). Study includes different turbulence modelling and mesh sensitivity analysis. Comparison with the WTT results is explored, and lessons learnt are collected.
Paryz, Roman W.
Several upgrade projects have been completed at the NASA Langley Research Center National Transonic Facility over the last 1.5 years in an effort defined as STARBUKS - Subsonic Transonic Applied Refinements By Using Key Strategies. This multi-year effort was undertaken to improve NTF's overall capabilities by addressing Accuracy and Validation, Productivity, and Reliability areas at the NTF. This presentation will give a brief synopsis of each of these efforts.
5. VIEW LOOKING NORTH AT 8-FOOT TRANSONIC PRESSURE TUNNEL PLENUM FLOOR AREA. NOTE SCHLIEREN OPTICAL SYSTEM ON STRUCTURE AT RIGHT CENTER. - NASA Langley Research Center, 8-Foot Transonic Pressure Tunnel, 640 Thornell Avenue, Hampton, Hampton, VA
1. VIEW LOOKING SOUTHEAST AT EXTERIOR OF 8-FOOT TRANSONIC PRESSURE TUNNEL. NOTE EXPANSION RINGS. - NASA Langley Research Center, 8-Foot Transonic Pressure Tunnel, 640 Thornell Avenue, Hampton, Hampton, VA
2. VIEW LOOKING EAST-NORTHEAST AT EXTERIOR OF 8-FOOT TRANSONIC PRESSURE TUNNEL (BUILDING 640). NOTE NACA LOGO OVER DOORWAY. - NASA Langley Research Center, 8-Foot Transonic Pressure Tunnel, 640 Thornell Avenue, Hampton, Hampton, VA
Tanaka, K.; Hirose, H.
The development of transonic aerodynamic computation methods and specific examples, as well as examples of three-dimensional transonic computation in design, are discussed. The case of the transonic transport and the case of the small transport are analyzed. Requirements for programs of the future are itemized.
Bodapati, S.; Lee, C.-S.
The steady and unsteady wake profiles of an airfoil with an oscillating flap were measured at nominal free stream Mach number of 0.8 in the NASA Ames 11 x 11-foot wind tunnel. The instantaneous wake velocity and pressure profiles at four axial locations are presented up to one chord length from the trailing edge. Both fundamental harmonic frequency and typical time history data are presented to observe the effects of airfoil incidence and flap angle. The drag coefficient obtained from the wake pressure measurements is compared with that obtained from the airfoil pressure distribution.
Batina, John T.; Seidel, David A.; Bland, Samuel R.; Bennett, Robert M.
A transonic unsteady aerodynamic and aeroelasticity code has been developed for application to realistic aircraft configurations. The new code is called CAP-TSD which is an acronym for Computational Aeroelasticity Program - Transonic Small Disturbance. The CAP-TSD code uses a time-accurate approximate factorization (AF) algorithm for solution of the unsteady transonic small-disturbance equation. The AF algorithm is very efficient for solution of steady and unsteady transonic flow problems. It can provide accurate solutions in only several hundred time steps yielding a significant computational cost savings when compared to alternative methods. The new code can treat complete aircraft geometries with multiple lifting surfaces and bodies including canard, wing, tail, control surfaces, launchers, pylons, fuselage, stores, and nacelles. Applications are presented for a series of five configurations of increasing complexity to demonstrate the wide range of geometrical applicability of CAP-TSD. These results are in good agreement with available experimental steady and unsteady pressure data. Calculations for the General Dynamics one-ninth scale F-16C aircraft model are presented to demonstrate application to a realistic configuration. Unsteady results for the entire F-16C aircraft undergoing a rigid pitching motion illustrated the capability required to perform transonic unsteady aerodynamic and aeroelastic analyses for such configurations.
Rivers, S. Melissa B.; Owens, Lewis R.; Wahls, Richard A.
Several computational studies were conducted as part of NASA s High Speed Research Program. Results of turbulence model comparisons from two studies on supersonic transport configurations performed during the NASA High-Speed Research program are given. The effects of grid topology and the representation of the actual wind tunnel model geometry are also investigated. Results are presented for both transonic conditions at Mach 0.90 and supersonic conditions at Mach 2.48. A feature of these two studies was the availability of higher Reynolds number wind tunnel data with which to compare the computational results. The transonic wind tunnel data was obtained in the National Transonic Facility at NASA Langley, and the supersonic data was obtained in the Boeing Polysonic Wind Tunnel. The computational data was acquired using a state of the art Navier-Stokes flow solver with a wide range of turbulence models implemented. The results show that the computed forces compare reasonably well with the experimental data, with the Baldwin-Lomax with Degani-Schiff modifications and the Baldwin-Barth models showing the best agreement for the transonic conditions and the Spalart-Allmaras model showing the best agreement for the supersonic conditions. The transonic results were more sensitive to the choice of turbulence model than were the supersonic results.
Burns, Ross A.; Danehy, Paul M.
Femtosecond laser electronic excitation tagging (FLEET) velocimetry was used to study the flowfield around a symmetric, transonic airfoil in the NASA Langley 0.3-m TCT facility. A nominal Mach number of 0.85 was investigated with a total pressure of 125 kPa and total temperature of 280 K. Two-components of velocity were measured along vertical profiles at different locations above, below, and aft of the airfoil at angles of attack of 0 deg, 3.5 deg, and 7deg. Measurements were assessed for their accuracy, precision, dynamic range, spatial resolution, and overall measurement uncertainty in the context of the applied flowfield. Measurement precisions as low as 1 m/s were observed, while overall uncertainties ranged from 4 to 5 percent. Velocity profiles within the wake showed sufficient accuracy, precision, and sensitivity to resolve both the mean and fluctuating velocities and general flow physics such as shear layer growth. Evidence of flow separation is found at high angles of attack.
Biollo, Roberto; Benini, Ernesto
Transonic axial flow compressors are fundamental components in aircraft engines as they make it possible to maximize pressure ratios per stage unit. This is achieved through a careful combination of both tangential flow deflections and, above all, by taking advantage of shock wave formation around the rotor blades. The resulting flow field is really complex as it features highly three-dimensional inviscid/viscous structures, strong shock-boundary layer interaction and intense tip clearance effects which negatively influence compressor efficiency. Complications are augmented at part load operation, where stall-related phenomena occur. Therefore, considerable research efforts are being spent, both numerically and experimentally, to improve efficiency and stall margin at peak efficiency and near stall operation. The present work aims at giving a complete review of the most recent advances in the field of aerodynamic design and operation of such machines. A great emphasis has been given to highlight the most relevant contribution in this field and to suggest the prospects for future developments.
The objective of the Subsonic/Transonic Stall Flutter Program was to obtain detailed measurements of both the steady and unsteady flow field surrounding a rotor and the mechanical state of the rotor while it was operating in both steady and flutter modes to provide a basis for future analysis and for development of theories describing the flutter phenomenon. The program revealed that while all blades flutter at the same frequency, they do not flutter at the same amplitude, and their interblade phase angles are not equal. Such a pattern represents the superposition of a number of rotating nodal diameter patterns, each characterized by a different amplitude and different phase indexing, but each rotating at a speed that results in the same flutter frequency as seen in the rotor system. Review of the steady pressure contours indicated that flutter may alter the blade passage pressure distribution. The unsteady pressure amplitude contour maps reveal regions of high unsteady pressure amplitudes near the leading edge, lower amplitudes near the trailing.
Campbell, Richard L.; Smith, Leigh A.
A method for designing wings and airfoils at transonic speeds using a predictor/corrector approach was developed. The procedure iterates between an aerodynamic code, which predicts the flow about a given geometry, and the design module, which compares the calculated and target pressure distributions and modifies the geometry using an algorithm that relates differences in pressure to a change in surface curvature. The modular nature of the design method makes it relatively simple to couple it to any analysis method. The iterative approach allows the design process and aerodynamic analysis to converge in parallel, significantly reducing the time required to reach a final design. Viscous and static aeroelastic effects can also be accounted for during the design or as a post-design correction. Results from several pilot design codes indicated that the method accurately reproduced pressure distributions as well as the coordinates of a given airfoil or wing by modifying an initial contour. The codes were applied to supercritical as well as conventional airfoils, forward- and aft-swept transport wings, and moderate-to-highly swept fighter wings. The design method was found to be robust and efficient, even for cases having fairly strong shocks.
Gartner, Jeremy; Amitay, Michael
In some airplanes such as fighter jets and UAV, short inlet ducts replace the more conventional ducts due to their shorter length. However, these ducts are associated with low length-to-diameter ratio and low aspect ratio and, thus, experience massive separation and the presence of secondary flow structures. These flow phenomena are undesirable as they lead to pressure losses and distortion at the Aerodynamic Interface Plane (AIP), where the engine face is located. It causes the engine to perform with a lower efficiency as it would with a straight duct diffuser. Different flow control techniques were studied on the short inlet duct, with the goal to reattach the flow and minimize the distortions at the AIP. Due to the complex interaction between the separation and the secondary flow structures, the necessity to understand the flow mechanisms, and how to control them at a more fundamental level, a new transonic diffuser with an upper ramp and a straight floor was designed and built. The objective of this project is to explore the effectiveness of different flow control techniques in a high subsonic (up to Mach 0.8) diffuser, so that the quasi two-dimensional separation and the formation of secondary flow structure can be isolated using a canonical flow field. Supported by Northrop Grumman.
Fasanella, Edwin L.; Robinson, Martha P.
The nonlinear dynamic finite element code DYnamic Crash Analysis of STructures (DYCAST) was used to model an accident scenario that occurred at the National Transonic Facility (NTF) wind tunnel. A post mishap investigation revealed that a total of five upstream bulkhead fairing plates were missing, three in one location and two in another. These plates were drawn into the wind tunnel's composite fan blades causing extensive damage. A DYCAST model was developed to determine if one-half of a small thermal shield flange clamp, weighing approximately 2.7 lbs., could have spun off the NTF drive shaft and impacted the bulkhead fairing plates with sufficient energy to cause failure of the attachment bolts. The clamp was presumed to have spun off at a tangent from the NTF drive shaft at a velocity of 1624 in/sec (drive shaft rotating at 580 rpm). The DYCAST analytical model predicts that impact of the 2.7 lbs projectile failed all of the bolts in two of the fairing plates allowing them to escape from the bulkhead ring with a low velocity of a few in/sec.
Lewis, M. C.
Validation data from the Transonic Self-Streamlining Wind Tunnel has proved the feasibility of streamlining two dimensional flexible walls at low speeds and up to transonic speeds, the upper limit being the speed where the flexible walls are just supercritical. At this condition, breakdown of the wall setting strategy is evident in that convergence is neither as rapid nor as stable as for lower speeds, and wall streamlining criteria are not always completely satisfied. The only major step necessary to permit the extension of two dimensional testing into higher transonic speeds is the provision of a rapid algorithm to solve for mixed flow in the imagery flow fields. The status of two dimensional high transonic testing in the Transonic Self-Streamlining Wind Tunnel is outlined and, in particular, the progress of adapting an algorithm, which solves the Transonic Small Perturbation Equation, for predicting the imagery flow fields is detailed.
A new small-disturbance model for combustion with transonic flow in a converging-diverging channel is presented. Attention is confined to dilute premixtures so that exothermicity is weak. The model explores the nonlinear interactions between the near-sonic speed of the flow, the small geometry changes from a straight channel, and the small heat release due to the chemical reaction. The asymptotic analysis results in the similarity parameters that govern the reacting flow field. Also, the problem can be described by a nonhomogeneous (extended) transonic small-disturbance (TSD) equation which is coupled with an ordinary differential equation for the calculation of the reactant mass fraction in the combustible gas. An iterative numerical scheme which combines the Murman and Cole (1971) method for the solution of the TSD equation with Simpson's integration rule for the estimation of the reactant mass fraction is developed. The model is used to study transonic combustion at various inlet temperatures and Mach numbers.
Rosen, Bruce S.
Development of a computational method for prediction of external store carriage characteristics at transonic speeds is described. The geometric flexibility required for treatment of pylon-mounted stores is achieved by computing finite difference solutions on a five-level embedded grid arrangement. A completely automated grid generation procedure facilitates applications. Store modeling capability consists of bodies of revolution with multiple fore and aft fins. A body-conforming grid improves the accuracy of the computed store body flow field. A nonlinear relaxation scheme developed specifically for modified transonic small disturbance flow equations enhances the method's numerical stability and accuracy. As a result, treatment of lower aspect ratio, more highly swept and tapered wings is possible. A limited supersonic freestream capability is also provided. Pressure, load distribution, and force/moment correlations show good agreement with experimental data for several test cases. A detailed computer program description for the Transonic Store Carriage Loads Prediction (TSCLP) Code is included.
Hafez, M. M.; Murman, E. M.; Wellford, L. C.
The paper reviews the chief finite difference and finite element techniques used for numerical solution of nonlinear mixed elliptic-hyperbolic equations governing transonic flow. The forms of the governing equations for unsteady two-dimensional transonic flow considered are the Euler equation, the full potential equation in both conservative and nonconservative form, the transonic small-disturbance equation in both conservative and nonconservative form, and the hodograph equations for the small-disturbance case and the full-potential case. Finite difference methods considered include time-dependent methods, relaxation methods, semidirect methods, and hybrid methods. Finite element methods include finite element Lax-Wendroff schemes, implicit Galerkin method, mixed variational principles, dual iterative procedures, optimal control methods and least squares.
A rapid computation of a sequence of transonic flow solutions has to be performed in many areas of aerodynamic technology. The employment of low-cost vector array processors makes the conduction of such calculations economically feasible. However, for a full utilization of the new hardware, the developed algorithms must take advantage of the special characteristics of the vector array processor. The present investigation has the objective to develop an efficient algorithm for solving transonic flow problems governed by mixed partial differential equations on an array processor.
Carlson, L. A.
An inverse numerical technique for designing transonic airfoils having a prescribed pressure distribution is presented. The method uses the full potential equation, inverse boundary conditions, and Cartesian coordinates. It includes simultaneous airfoil update and utilizes a direct-inverse approach that permits a logical method for controlling trailing edge closure. The method can also be used for the analysis of flowfields about specified airfoils. Comparison with previous results shows that accurate results can be obtained with a Cartesian grid. Examples show the application of the method to design aft-cambered and other airfoils specifically for transonic flight.
Rivera, Jose A.; Florance, James R.
The Transonic Dynamics Tunnel (TDT) became in operational in 1960, and since that time has achieved the status of the world's premier wind tunnel for testing large in aeroelastically scaled models at transonic speeds. The facility has many features that contribute to its uniqueness for aeroelastic testing. This paper will briefly describe these capabilities and features, and their relevance to aeroelastic testing. Contributions to specific airplane configurations and highlights from the flutter tests performed in the TDT aimed at investigating the aeroelastic characteristics of these configurations are presented.
Dowell, E. H.; Thomas, J. P.; Hall, K. C.
Limit cycle oscillations have been observed in flight operations of modern aircraft, wind tunnel experiments and mathematical models. Both fluid and structural nonlinearities are thought to contribute to these phenomena. With recent advances in reduced order aerodynamic modeling, it is now feasible to analyze limit cycle oscillations that may occur in transonic flow including the effects of structural and fluid nonlinearities. In this paper an airfoil with control surface freeplay (a common structural nonlinearity) is used to investigate transonic flutter and limit cycle oscillations. The reduced order aerodynamic model used in this paper assumes the shock motion is small and in proportion to the structural motions.
Foster, Jean M.; Adcock, Jerry B.
The National Transonic Facility is a complex cryogenic wind tunnel facility. This report briefly describes the facility, the data systems, and the instrumentation used to acquire research data. The computational methods and equations are discussed in detail and many references are listed for those who need additional technical information. This report is intended to be a user's guide, not a programmer's guide; therefore, the data reduction code itself is not documented. The purpose of this report is to assist personnel involved in conducting a test in the National Transonic Facility.
Kilgore, W. Allen; Chan, David; Balakrishna, S.; Wahls, Richard A.
The U.S. National Transonic Facility continues to make enhancements to provide quality data in a safe, efficient and cost effective method for aerodynamic ground testing. Recent enhancements discussed in this paper include the development of a Mixed-mode of operations that combine Air-mode operations with Nitrogen-mode operations. This implementation and operational results of this new Mixed-mode expands the ambient temperature transonic region of testing beyond the Air-mode limitations at a significantly reduced cost over Nitrogen Mode operation.
Chen, Zuoyi; Guo, Jingrong
The use of the Hodograph Method in the design of a transonic compressor cascade is discussed. The design of the flow mode in the transonic compressor cascade must be as follows: the flow in the nozzle part should be uniform and smooth; the location of the sonic line should be reasonable; and the aerodynamic character of the flow canal in the subsonic region should be met. The rate through cascade may be determined by the velocity distribution in the subsonic region (i.e., by the numerical solution of the Chaplygin equation). The supersonic sections A'C' and AD are determined by the analytical solution of the Mixed-Type Hodograph equation.
This report is an extension of a previous investigation (described in NACA rep. 1069) concerned with the solution of the nonlinear differential equation for transonic flow past a wavy wall. In the present work several new notions are introduced which permit the solution of the recursion formulas arising from the method of integration in series. In addition, a novel numerical tests of convergence, applied to the power series (in transonic similarity parameter) representing the local Mach number distribution at the boundary, indicates that smooth symmetrical potential flow past the wavy wall is no longer possible once the critical value of the stream Mach number has been exceeded.
Liepmann, H. W.; Black, R. E.; Dietz, R. O.; Kirchner, M. E.; Sears, W. R.
The proposed National Transonic Facility (NTF) operational plan is reviewed. The NTF will provide an aerodynamic test capability significantly exceeding that of other transonic regime wind tunnels now available. A limited number of academic research program that might use the NTF are suggested. It is concluded that the NTF operational plan is useful for management, technical, instrumentation, and model building techniques available in the specialized field of aerodynamic analysis and simulation. It is also suggested that NASA hold an annual conference to discuss wind tunnel research results and to report on developments that will further improve the utilization and cost effectiveness of the NTF and other wind tunnels.
Issac, Jason Cherion; Kapania, Rakesh K.
Flutter analysis of a two degree of freedom airfoil in compressible flow is performed using a state-space representation of the unsteady aerodynamic behavior. Indicial response functions are used to represent the normal force and moment response of the airfoil. The structural equations of motion of the airfoil with bending and torsional degrees of freedom are coupled to the unsteady air loads and the aeroelastic system so modelled is solved as an eigenvalue problem to determine the stability. The aeroelastic equations are also directly integrated with respect to time and the time-domain results compared with the results from the eigenanalysis. A good agreement is obtained. The derivatives of the flutter speed obtained from the eigenanalysis are calculated with respect to the mass and stiffness parameters by both analytical and finite-difference methods for various transonic Mach numbers. The experience gained from the two degree of freedom model is applied to study the sensitivity of the flutter response of a wing with respect to various shape parameters. The parameters being considered are as follows: (1) aspect ratio; (2) surface area of the wing; (3) taper ratio; and (4) sweep. The wing deflections are represented by Chebyshev polynomials. The compressible aerodynamic state-space model used for the airfoil section is extended to represent the unsteady aerodynamic forces on a generally laminated tapered skewed wing. The aeroelastic equations are solved as an eigenvalue problem to determine the flutter speed of the wing. The derivatives of the flutter speed with respect to the shape parameters are calculated by both analytical and finite difference methods.
Wolf, S. W. D.
Twenty four test runs of the Transonic Self-Streamlining Wind Tunnel were performed with the flexible walls 'streamlined' around a two dimensional section of four inch chord, over the Mach number range 0.3 to 0.89. Relevant wall and model data for the streamlined cases are presented.
TERMS Transonic, Compressor, Pressure Instability, Low Dominant Frequencies, Turbomachinery , Near Stall Disturbances. 15. NUMBER OF PAGES 153 16...pre-stall operations. Due to technological advancement in turbomachinery , disturbances that have potential to lead to instability within a...This technique fills a gap in current turbomachinery related literature because it provides a better understanding of modal disturbances in
Carlson, L. A.
A program called TRANDES is presented that is used for the analysis of steady, irrotational transonic flow over specified two-dimensional airfoils in free air or for the design of airfoils having a prescribed pressure distribution, including the effects of weak viscous interaction. Instructions on program usage, listings of the program, and sample cases are given.
Lepicovsky, J.; McFarland, E. R.; Chima, R. V.; Capece, V. R.; Hayden, J.
A study was conducted in the NASA Glenn Research Center linear cascade on the intermittent flow on the suction surface of an airfoil section from the tip region of a modern low aspect ratio fan blade. Experimental results revealed that, at a large incidence angle, a range of transonic inlet Mach numbers exist where the leading-edge shock-wave pattern was unstable. Flush mounted high frequency response pressure transducers indicated large local jumps in the pressure in the leading edge area, which generates large intermittent loading on the blade leading edge. These measurements suggest that for an inlet Mach number between 0.9 and 1.0 the flow is bi-stable, randomly switching between subsonic and supersonic flows. Hence, it appears that the change in overall flow conditions in the transonic region is based on the frequency of switching between two stable flow states rather than on the continuous increase of the flow velocity. To date, this flow behavior has only been observed in a linear transonic cascade. Further research is necessary to confirm this phenomenon occurs in actual transonic fans and is not the byproduct of an endwall restricted linear cascade.
Li, Wesley Waisang; Pak, Chan-Gi
A technique for approximating the modal aerodynamic influence coefficients [AIC] matrices by using basis functions has been developed and validated. An application of the resulting approximated modal AIC matrix for a flutter analysis in transonic speed regime has been demonstrated. This methodology can be applied to the unsteady subsonic, transonic and supersonic aerodynamics. The method requires the unsteady aerodynamics in frequency-domain. The flutter solution can be found by the classic methods, such as rational function approximation, k, p-k, p, root-locus et cetera. The unsteady aeroelastic analysis for design optimization using unsteady transonic aerodynamic approximation is being demonstrated using the ZAERO(TradeMark) flutter solver (ZONA Technology Incorporated, Scottsdale, Arizona). The technique presented has been shown to offer consistent flutter speed prediction on an aerostructures test wing [ATW] 2 configuration with negligible loss in precision in transonic speed regime. These results may have practical significance in the analysis of aircraft aeroelastic calculation and could lead to a more efficient design optimization cycle
Li, Wesley W.; Pak, Chan-gi
A technique for approximating the modal aerodynamic influence coefficients matrices by using basis functions has been developed and validated. An application of the resulting approximated modal aerodynamic influence coefficients matrix for a flutter analysis in transonic speed regime has been demonstrated. This methodology can be applied to the unsteady subsonic, transonic, and supersonic aerodynamics. The method requires the unsteady aerodynamics in frequency-domain. The flutter solution can be found by the classic methods, such as rational function approximation, k, p-k, p, root-locus et cetera. The unsteady aeroelastic analysis for design optimization using unsteady transonic aerodynamic approximation is being demonstrated using the ZAERO flutter solver (ZONA Technology Incorporated, Scottsdale, Arizona). The technique presented has been shown to offer consistent flutter speed prediction on an aerostructures test wing 2 configuration with negligible loss in precision in transonic speed regime. These results may have practical significance in the analysis of aircraft aeroelastic calculation and could lead to a more efficient design optimization cycle.
Batina, John T.
Improvements to a time-accurate approximate factorization (AF) algorithm were implemented for steady and unsteady transonic analysis of realistic aircraft configurations. These algorithm improvements were made to the CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code developed at the Langley Research Center. The code permits the aeroelastic analysis of complete aircraft in the flutter critical transonic speed range. The AF algorithm of the CAP-TSD code solves the unsteady transonic small-disturbance equation. The algorithm improvements include: an Engquist-Osher (E-O) type-dependent switch to more accurately and efficiently treat regions of supersonic flow; extension of the E-O switch for second-order spatial accuracy in these regions; nonreflecting far field boundary conditions for more accurate unsteady applications; and several modifications which accelerate convergence to steady-state. Calculations are presented for several configurations including the General Dynamics one-ninth scale F-16C aircraft model to evaluate the algorithm modifications. The modifications have significantly improved the stability of the AF algorithm and hence the reliability of the CAP-TSD code in general.
Airloads investigation of an 0.030-scale model of the space shuttle vehicle 140A/B launch configuration (model 47-OTS) in the arc 11-foot unitary plan wind tunnel for Mach range 0.6 to 1.4 (IA14A), Volume 2
Gillins, R. L.
Results of tests conducted on an 0.030-scale launch configuration model of the space shuttle vehicle 140A/B in the NASA/ARC 11-foot unitary plan wind tunnel are presented. Aerodynamic loads data were obtained at Mach numbers from 0.6 to 1.4. Surface pressure distributions were obtained simultaneously with six-component stability and control force data on the complete launch configuration. The configuration consisted of the orbiter, an external tank, two solid rocket boosters, and associated intercomponent attach hardware. Angles of attack and sideslip from -10 degrees to +10 degrees were investigated.
Harder, Keith C; Klunker, E B
The basic ideas of the slender-body approximation have been applied to the nonlinear transonic-flow equation for the velocity potential in order to obtain some of the essential features of slender-body theory at transonic speeds. The results of the investigation are presented from a unified point of view which demonstrates the similarity of slender-body solutions in the various Mach number ranges. The transonic area rule and some conditions concerning its validity follow from the analysis.
Newman, P. A.; Kemp, W. B., Jr.; Garriz, J. A.
Several nonlinear transonic codes and a panel method code for wind tunnel/wall interference assessment and correction (WIAC) studies are reviewed. Contrasts between two- and three-dimensional transonic testing factors which affect WIAC procedures are illustrated with airfoil data from the NASA/Langley 0.3-meter transonic cyrogenic tunnel and Pathfinder I data. Also, three-dimensional transonic WIAC results for Mach number and angle-of-attack corrections to data from a relatively large 20 deg swept semispan wing in the solid wall NASA/Ames high Reynolds number Channel I are verified by three-dimensional thin-layer Navier-Stokes free-air solutions.
Hafez, M. M.; Murman, E. M.
The prediction of aerodynamic forces in the transonic regime generally requires a flow field calculation to solve the governing non-linear mixed elliptic-hyperbolic partial differential equations. Finite difference techniques were developed to the point that design and analysis application are routine, and continual improvements are being made by various research groups. The principal limitation in extending finite difference methods to complex three-dimensional geometries is the construction of a suitable mesh system. Finite element techniques are attractive since their application to other problems have permitted irregular mesh elements to be employed. The purpose of this paper is to review the recent developments in the application of finite element methods to transonic flow problems and to report some recent results.
Stainback, P. C.
The present paper provides a review of hot wire anemometry for compressible flows, giving particular attention to the transonic flow problem. It is pointed out that the first and most important definitive work in hot wire anemometry for compressible flows was reported by Kovasznay (1953). The existence of three independent fluctuating modes in compressible flows for small perturbations was found, taking into account the vorticity mode, the entropy mode, and the sound-wave mode. A review of Kovasznays' method for supersonic flows is also presented, and advances reported by Markovin (1956) are examined. Attention is given to experiments conducted by Horstman and Rose (1977), a general solution to the hot wire problem at transonic conditions sought by Stainback et al. (1983), and some apparent problems.
Vinh, Lam-Son; Edwards, John W.; Seidel, David A.; Batina, John T.
Implementation of a capability to calculate longitudinal short-period response in the CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) finite-difference code is described. The code, developed recently at the NASA Langley Research Center, is capable of solving steady and unsteady flows about complete aircraft configurations and is used primarily for aeroelastic calculations in the critical transonic speed range. The longitudinal short-period equations of motion in state-space form have been coupled to the time-accurate lift and moment calculated by the program. Transient responses to an elevator pulse for free-flying aircraft demonstrate the new capability. A trim routine is also added to the code to obtain trim automatically during steady-state flow field convergence. Stability and control derivatives are estimated from the calculated transient response by a maximum likelihood estimation program. Results for a fighter configuration and a general aviation configuration are presented to assess the capability.
Durham, Michael H.; Cole, Stanley R.; Cazier, F. W., Jr.; Keller, Donald F.; Parker, Ellen C.; Wilkie, W. Keats
The flutter characteristics of a generic arrow-wing supersonic transport configuration are studied. The wing configuration has a 3 percent biconvex airfoil and a leading-edge sweep of 73 deg out to a cranked tip with a 60 deg leading-edge sweep. The ground vibration tests and flutter test procedure are described. The effects of flutter on engine nacelles, fuel loading, wing-mounted vertical fin, wing angle-of-attack, and wing tip mass and stiffness distributions are analyzed. The data reveal that engine nacelles reduce the transonic flutter dynamic pressure by 25-30 percent; fuel loadings decrease dynamic pressures by 25 percent; 4-6 deg wing angles-of-attack cause steep transonic boundaries; and 5-10 percent changes in flutter dynamic pressures are the result of the wing-mounted vertical fin and wing-tip mass and stiffness distributions.
Young, Clarence P., Jr.; Popernack, Thomas G., Jr.; Gloss, Blair B.
Vibrations of models and model support system were encountered during testing in the National Transonic Facility. Model support system yaw plane vibrations have resulted in model strain gage balance design load limits being reached. These high levels of vibrations resulted in limited aerodynamic testing for several wind tunnel models. The yaw vibration problem was the subject of an intensive experimental and analytical investigation which identified the primary source of the yaw excitation and resulted in attenuation of the yaw oscillations to acceptable levels. This paper presents the principal results of analyses and experimental investigation of the yaw plane vibration problems. Also, an overview of plans for development and installation of a permanent model system dynamic and aeroelastic response measurement and monitoring system for the National Transonic Facility is presented.
Isom, M. P.
Differential equations and boundary conditions for a rotor blade in forward flight, with subsonic or transonic tip Mach number, are derived. A variety of limiting flow regimes determined by different limits involving blade thickness ratio, aspect ratio, advance ratio and maximum tip Mach number is discussed. The transonic problem is discussed in some detail, and in particular the conditions that make this problem quasi-steady or essentially unsteady are determined. Asymptotic forms of equations and boundary conditions that are valid in an appropriately scaled region of the tip and an azimuthal sector on the advancing side are derived. The equations are then put in a form that is valid from the blade tip inboard through the strip theory region.
Sekula, Martin K.; Piatak, David J.; Rausch, Russ D.; Florance, James R.; Ramey, James M.
A series of wind tunnel tests were conducted at the NASA Langley Research Center Transonic Dynamics Tunnel to assess the transonic buffet environment for the Space Launch System (SLS) launch vehicle. An initial test, conducted in 2012, indicated an elevated buffet environment prompting a second test to provide further insight into the buffet phenomena and assess potential solutions to reduce the response levels of these environments. During the course of the test program, eight variants of the SLS-10000 configuration were examined. The effect of these configuration variants on the coefficient of the root-mean-square fluctuation of pressure about the mean as a function of test condition indicates that the maximum fluctuating pressure levels are extremely sensitive to the geometry of the forward attachment of the solid rocket boosters (SRBs) to the SLS Core. The addition of flow fences or changes to the SRB nose cone geometry can alleviate the unsteady pressure environment.
Crouch, Jeffrey D.; Sutanto, Mary I.; Witkowski, David P.; Watkins, A. Neal; Rivers, Melissa B.; Campbell, Richard L.
A transonic wing, designed to accentuate key transition physics, is tested at cryogenic conditions at the National Transonic Facility at NASA Langley. The collaborative test between Boeing and NASA is aimed at assessing the facility for high-Reynolds number testing of configurations with significant regions of laminar flow. The test shows a unit Reynolds number upper limit of 26 M/ft for achieving natural transition. At higher Reynolds numbers turbulent wedges emanating from the leading edge bypass the natural transition process and destroy the laminar flow. At lower Reynolds numbers, the transition location is well correlated with the Tollmien-Schlichting-wave N-factor. The low-Reynolds number results suggest that the flow quality is acceptable for laminar flow testing if the loss of laminar flow due to bypass transition can be avoided.
Smith, Leigh Ann; Campbell, Richard L.
Methodology was developed for designing airfoils and wings at transonic speeds which includes a technique that can account for static aeroelastic deflections. This procedure is capable of designing either supercritical or more conventional airfoil sections. Methods for including viscous effects are also illustrated and are shown to give accurate results. The methodology developed is an interactive system containing three major parts. A design module was developed which modifies airfoil sections to achieve a desired pressure distribution. This design module works in conjunction with an aerodynamic analysis module, which for this study is a small perturbation transonic flow code. Additionally, an aeroelastic module is included which determines the wing deformation due to the calculated aerodynamic loads. Because of the modular nature of the method, it can be easily coupled with any aerodynamic analysis code.
Walker, Eric L.
A validation test has recently been constructed for wall interference methods as applied to the National Transonic Facility (NTF). The goal of this study was to begin to address the uncertainty of wall-induced-blockage interference corrections, which will make it possible to address the overall quality of data generated by the facility. The validation test itself is not specific to any particular modeling. For this present effort, the Transonic Wall Interference Correction System (TWICS) as implemented at the NTF is the mathematical model being tested. TWICS uses linear, potential boundary conditions that must first be calibrated. These boundary conditions include three different classical, linear. homogeneous forms that have been historically used to approximate the physical behavior of longitudinally slotted test section walls. Results of the application of the calibrated wall boundary conditions are discussed in the context of the validation test.
Barnwell, R. W.
An analytical solution was obtained for the perturbation velocity potential for transonic flow about lifting wing-body configurations with order-one span-length ratios and small reduced-span-length ratios and equivalent-thickness-length ratios. The analysis is performed with the method of matched asymptotic expansions. The angles of attack which are considered are small but are large enough to insure that the effects of lift in the region far from the configuration are either dominant or comparable with the effects of thickness. The modification to the equivalence rule which accounts for these lift effects is determined. An analysis of transonic flow about lifting wings with large aspect ratios is also presented.
What may appear at first glance to be a swimming shark is a wind tunnel model of the Space Shuttle Orbiter, being tested at NASA's Langley Research Center in Hampton,VA. The Orbiter model is 5.5 feet long (1/20th of the real Orbiter's length) and has remotely operated control surfaces. Inside Langley's 16 foot Transonic Wind Tunnel, the model simulated Orbiter re-entry into the Earth's atmosphere, when it must fly through the transonic speed range (the range that crosses the sound barrier). Information on Orbiter stability and control, collected and analyzed during the tests, were integrated with other data to become part of computerized flight simulation programs.
Wahls, R. N.; Owens, L. R.; Rivers, S. M. B.
A High Speed Civil Transport configuration was tested in the National Transonic Facility at the NASA Langley Research Center as part of NASA's High Speed Research Program. The primary purposes of the tests were to assess Reynolds number scale effects and the high Reynolds number aerodynamic characteristics of a realistic, second generation supersonic transport while providing data for the assessment of computational methods. The tests included longitudinal and lateral/directional studies at low speed high-lift and transonic conditions across a range of Reynolds numbers from that available in conventional wind tunnels to near flight conditions. Results are presented which focus on both the Reynolds number and static aeroelastic sensitivities of longitudinal characteristics at Mach 0.90 for a configuration without an empennage.
Gibbons, Michael D.
A flexible fuselage capability has been developed and implemented within version 1.2 of the CAP-TSD code. The capability required adding time dependent terms to the fuselage surface boundary conditions and the fuselage surface pressure coefficient. The new capability will allow modeling the effect of a flexible fuselage on the aeroelastic stability of complex configurations. To assess the flexible fuselage capability several steady and unsteady calculations have been performed for slender fuselages with circular cross-sections. Steady surface pressures are compared with experiment at transonic flight conditions. Unsteady cross-sectional lift is compared with other analytical results at a low subsonic speed and a transonic case has been computed. The comparisons demonstrate the accuracy of the flexible fuselage modifications.
Foughner, J. T., Jr.; Reed, J. F.; Wynne, E. C.
Wind-tunnel tests have been made in the Langley transonic dynamics tunnel on a 0.25-scale model of Sandia Laboratories' 3.96-meter (13-foot), slanted ribbon design, lifting parachute. The lifting parachute is the first stage of a proposed two-stage payload delivery system. The lifting parachute model was attached to a forebody representing the payload. The forebody was designed and installed in the test section in a manner which allowed rotational freedom about the pitch and yaw axes. Values of parachute axial force coefficient, rolling moment coefficient, and payload trim angles in pitch and yaw are presented through the transonic speed range. Data are presented for the parachute in both the reefed and full open conditions. Time history records of lifting parachute deployment and disreefing tests are included.
Lyrintzis, A. S.
Several parameters of transonic blade-vortex interactions (BVI) are being studied and some ideas for noise reduction are introduced and tested using numerical simulation. The model used is the two-dimensional high frequency transonic small disturbance equation with regions of distributed vorticity (VTRAN2 code). The far-field noise signals are obtained by using the Kirchhoff method with extends the numerical 2-D near-field aerodynamic results to the linear acoustic 3-D far-field. The BVI noise mechanisms are explained and the effects of vortex type and strength, and angle of attack are studied. Particularly, airfoil shape modifications which lead to noise reduction are investigated. The results presented are expected to be helpful for better understanding of the nature of the BVI noise and better blade design.
Computer aided tomography (CAT) provides a means of noninvasively measuring the air density distribution around an aerodynamic model. This technique is global in that a large portion of the flow field can be measured. A test of the applicability of CAT to transonic velocities was studied. A hemispherical-nose cylinder afterbody model was tested at a Mach number of 0.8 with a new laser holographic interferometer at the 2- by 2-Foot Transonic Wind Tunnel. Holograms of the flow field were taken and were reconstructed into interferograms. The fringe distribution (a measure of the local densities) was digitized for subsequent data reduction. A computer program based on the Fourier-transform technique was developed to convert the fringe distribution into three-dimensional densities around the model. Theoretical aerodynamic densities were calculated for evaluating and assessing the accuracy of the data obtained from the tomographic method.
Kilgore, W. A.; Balakrishna, S.; Butler, D. H.
This paper describes the results of recent efforts to reduce the tunnel dynamics at the National Transonic Facility. The results presented describe the findings of an extensive data analysis, the proposed solutions to reduce dynamics and the results of implementing these solutions. These results show a 90% reduction in the dynamics around the model support structure and a small impact on reducing model dynamics. Also presented are several continuing efforts to further reduce dynamics.
Klich, P. J.; Richards, W. H.; Ahl, E. L., Jr.
The test program for the basic prepreg materials used in process development work and planned fabrication of the national transonic facility fan blade is presented. The basic prepreg materials and the design laminate are characterized at 89 K, room temperature, and 366 K. Characterization tests, test equipment, and test data are discussed. Material tests results in the warp direction are given for tensile, compressive, fatigue (tension-tension), interlaminar shear and thermal expansion.
Whitlow, W., Jr.
Characteristic far-field boundary conditions for the three-dimensional unsteady transonic small disturbance potential equation have been developed. The boundary conditions were implemented in the XTRAN3S finite difference code and tested for a flat plate rectangular wing with a pulse in angle of attack; the freestream Mach number was 0.85. The calculated force response shows that the characteristic boundary conditions reduce disturbances that are reflected from the computational boundaries.
Stewart, Pamela N.
The data quality analysis program that was developed at Langley Research Center's National Transonic Facility is described. The program provides a computer driven systematic analysis of data taken during calibrations of the high speed digital data acquisition system. Five distinct checks that are performed on the calibration data are outlined. The five checks are for non-linearity, noise, short term drift, long term drift, and the proper functioning of the calibrator. The program has established a standard set of evaluation guidelines.
An extended integral equation method for transonic flows is developed. In the extended integral equation method velocities in the flow field are calculated in addition to values on the aerofoil surface, in contrast with the less accurate 'standard' integral equation method in which only surface velocities are calculated. The results obtained for aerofoils in subcritical flow and in supercritical flow when shock waves are present compare satisfactorily with the results of recent finite difference methods.
Parker, P. A.
This paper provides an overview of force measurement at the National Transonic Facility (NTF). The NTF has unique force measurement requirements that dictate an integration of all aspects of balance design, production, and calibration. An overview of current force measurement capabilities is provided along with new balance development efforts. Research activities in the areas of thermal compensation and balance calibration are presented. Also, areas of future research are detailed.
Saaris, G. R.; Gilkey, R. D.; Smit, K. L.; Tinoco, E. N.
The application of a three-dimensional transonic flow analysis method, TRANAIR, is explored from the point of view of a user. Detailed features of the program are outlined to give a better understanding of capability. Numerous results are presented to show some of the complex configurations which have been analyzed. In particular, examples are provided which show the application to turbofan engine installation on transport aircraft.
Ballhaus, W. F.; Magnus, R.; Yoshihara, H.
A finite difference flutter analysis is presented for the NACA 64A-410 airfoil at M equals 0.72, where the incidence is abruptly changed from 2 to 4 degrees. The effect of gust loads is studied, and the unsteady flow adjusting process is displayed. The semi-implicit procedure of Ballhaus and Lomax (1974) is used to solve the small disturbance transonic potential equation. The physical aspects of the results, rather than the numerical details, are emphasized.
Steger, Joseph L.; Van Dalsem, William R.
Some of the basic finite-difference schemes that can be used to solve the nonlinear equations that describe unsteady inviscid and viscous transonic flow are reviewed. Numerical schemes for solving the unsteady Euler and Navier-Stokes, boundary-layer, and nonlinear potential equations are described. Emphasis is given to the elementary ideas used in constructing various numerical procedures, not specific details of any one procedure.
Fig. 20 shows the instantaneous pressures on an NACA 0012 airfoil oscillating in pitch about its quarter chord. In this case, M = 0.755, a(t...Unsteady Transonic Small Disturbance Equation. NASA TM 85723, 1983. Landon, R. H.: NACA 0012. Oscillatory and Transient Pitching. Compendium of...of aspect-ratio 6 rectangular wing with NACA 0012 airfoil at Mach lumber 0.82, 0=0 1-14 0 o Integral equation 0 Experinnent r Sonic
Edwards, John W.; Schuster, David M.; Spain, Charles V.; Keller, Donald F.; Moses, Robert W.
The Models for Aeroelastic Validation Research Involving Computation semi-span wind-tunnel model (MAVRIC-I), a business jet wing-fuselage flutter model, was tested in NASA Langley's Transonic Dynamics Tunnel with the goal of obtaining experimental data suitable for Computational Aeroelasticity code validation at transonic separation onset conditions. This research model is notable for its inexpensive construction and instrumentation installation procedures. Unsteady pressures and wing responses were obtained for three wingtip configurations of clean, tipstore, and winglet. Traditional flutter boundaries were measured over the range of M = 0.6 to 0.9 and maps of Limit Cycle Oscillation (LCO) behavior were made in the range of M = 0.85 to 0.95. Effects of dynamic pressure and angle-of-attack were measured. Testing in both R134a heavy gas and air provided unique data on Reynolds number, transition effects, and the effect of speed of sound on LCO behavior. The data set provides excellent code validation test cases for the important class of flow conditions involving shock-induced transonic flow separation onset at low wing angles, including LCO behavior.
Schmeer, James W.; Salters, Leland B., Jr.; Cassetti, Marlowe D.
An investigation of the transonic performance characteristics of several noise-suppressor configurations has been conducted in the Langley 16-foot transonic tunnel. The models were tested statically and over a Mach number range from 0.70 to 1.05 at an angle of attack of 0 deg. The primary jet total-pressure ratio was varied from 1.0 (jet off) to about 4.5. The effect of secondary air flow on the performance of two of the configurations was investigated. A hydrogen peroxide turbojet-engine simulator was used to supply the hot-jet exhaust. An 8-lobe afterbody with centerbody, short shroud, and secondary air had the highest thrust-minus-drag coefficients of the six noise-suppressor configurations tested. The 12-tube and 12-lobe afterbodies had the lowest internal losses. The presence of an ejector shroud partially shields the external pressure distribution of the 8-lobe after-body from the influence of the primary jet. A ring-airfoil shroud increased the static thrust of the annular nozzle but generally decreased the thrust minus drag at transonic Mach numbers.
Vincenti, Walter G; Wagoner, Cleo B
A theoretical study has been made of the aerodynamic characteristics at zero angle of attack of a thin, doubly symmetrical double-wedge profile in the range of supersonic flight speed in which the bow wave is detached. The analysis utilizes the equations of the transonic small-disturbance theory and involves no assumptions beyond those implicit in this theory. The mixed flow about the front half of the profile is calculated by relaxation solution of boundary conditions along the shock polar and sonic line. The purely subsonic flow about the rear of the profile is found by means of the method of characteristics specialized to the transonic small-disturbance theory. Complete calculations were made for four values of the transonic similarity parameter. These were found sufficient to bridge the gap between the previous results of Guderley and Yoshihara at a Mach number of 1 and the results which are readily obtained when the bow wave is attached and the flow is completely supersonic.
Schlecht, Robin; Anders, Scott
Wind-tunnel tests were conducted in the NASA Langley Transonic Dynamics Tunnel on a 6 percent-thick, elliptical circulation-control airfoil with upper-surface and lower-surface blowing capability. Results for elliptical Coanda trailing-edge geometries, biconvex Coanda trailing-edge geometries, and leading-edge geometries are reported. Results are presented at subsonic and transonic Mach numbers of 0.3 and 0.8, respectively. When considering one fixed trailing-edge geometry, for both the subsonic and transonic conditions it was found that the [3.0:1] ratio elliptical Coanda surface with the most rounded leading-edge  performed favorably and was determined to be the best compromise between comparable configurations that took advantage of the Coanda effect. This configuration generated a maximum. (Delta)C(sub 1) = 0.625 at a C(sub mu) = 0.06 at M = 0.3, alpha = 6deg. This same configuration generated a maximum (Delta)C(sub 1) = 0.275 at a C(sub mu) = 0.0085 at M = 0.8, alpha = 3deg.
Reid, L.; Urasek, D. C.
A single-stage transonic compressor was tested with two rotor blade leading-edge configurations to investigate the effect of increased leading-edge thickness on the performance of a transonic blade row. The original rotor blade configuration was modified by cutting back the leading edge sufficiently to double the blade leading-edge thickness and thus the blade gap blockage in the tip region. At design speed this modification resulted in a decrease in rotor overall peak efficiency of four points. The major portion of this decrement in rotor overall peak efficienty was attributed to the flow conditions in the outer 30 percent of the blade span. At 70 and 90 percent of design speed, the modification had very little effect on rotor overall performance.
Kilgore, R. A.
A fan-driven transonic cryogenic tunnel was designed, and its purging, cooldown, and warmup times were determined satisfactory. Cooling with liquid nitrogen is at the power levels required for transonic testing. Good temperature distributions are obtained by using a simple nitrogen injection system.
improvement investigations to the Naval Postgraduate School Turbo Propulsion Laboratory’s (NPS TPL) Transonic Axially Splittered Rotor were investigated ...School Turbo Propulsion Laboratory’s (NPS TPL) Transonic Axially Splittered Rotor were investigated . Implementation of current NPS TPL design...16 Figure 12. Rotor downstream face and new mounting flange ..........................................17 Figure 13. Redesigned rotor shaft
4. VIEW LOOKING NORTH-NORTHEAST AT TEST SECTION OF 8-FOOT TRANSONIC PRESSURE TUNNEL SHOWING ACCESS PORT TO TEST SECTION (RIGHT) AND PLENUM SURROUNDING AREA. - NASA Langley Research Center, 8-Foot Transonic Pressure Tunnel, 640 Thornell Avenue, Hampton, Hampton, VA
The primary focus of this work is efficient aerodynamic shape optimization in transonic flow. Adjoint-based optimization techniques are employed on airfoil sections and evaluated in terms of computational accuracy as well as efficiency. This study examines two test cases proposed by the AIAA Aerodynamic Design Optimization Discussion Group. The first is a two-dimensional, transonic, inviscid, non-lifting optimization of a Modified-NACA 0012 airfoil. The second is a two-dimensional, transonic, viscous optimization problem using a RAE 2822 airfoil. The FUN3D CFD code of NASA Langley Research Center is used as the ow solver for the gradient-based optimization cases. Two shape parameterization techniques are employed to study their effect and the number of design variables on the final optimized shape: Multidisciplinary Aerodynamic-Structural Shape Optimization Using Deformation (MASSOUD) and the BandAids free-form deformation technique. For the two airfoil cases, angle of attack is treated as a global design variable. The thickness and camber distributions are the local design variables for MASSOUD, and selected airfoil surface grid points are the local design variables for BandAids. Using the MASSOUD technique, a drag reduction of 72.14% is achieved for the NACA 0012 case, reducing the total number of drag counts from 473.91 to 130.59. Employing the BandAids technique yields a 78.67% drag reduction, from 473.91 to 99.98. The RAE 2822 case exhibited a drag reduction from 217.79 to 132.79 counts, a 39.05% decrease using BandAids.
Bailey, F. R.
A relaxation method is described for the numerical solution of the transonic small disturbance equation for flow about a slender body of revolution. Results for parabolic arc bodies, both with and without an attached sting, are compared with wind-tunnel measurements for a free-stream Mach number range from 0.90 to 1.20. The method is also used to show the effects of wind-tunnel wall interference by including boundary conditions representing porous-wall and open-jet wind-tunnel test sections.
Boppe, C. W.; Stern, M. A.
A computational method which simulates transonic flow about wing-fuselage configurations has been extended to include the treatment of multiple body and non-planar wing surfaces. The finite difference relaxation scheme is characterized by a modified small disturbance flow equation and multiple embedded grid system. Wing-body combinations with as many as four nacelles/pods, four pylons, and wing-tip-mounted winglets can be analyzed. A scheme for modeling inlet spillage and engine exhaust interference effects has been included. Computed results are correlated with experimental data for three transport configurations.
Mohler, Stanley R., Jr.
The Wind-US Computational Fluid Dynamics flow solver computed flow solutions for a transonic diffusing duct. The calculations used an unstructured (hexahedral) grid. The Spalart-Allmaras turbulence model was used. Static pressures along the upper and lower wall agreed well with experiment, as did velocity profiles. The effect of the smoothing input parameters on convergence and solution accuracy was investigated. The meaning and proper use of these parameters are discussed for the benefit of Wind-US users. Finally, the unstructured solver is compared to the structured solver in terms of run times and solution accuracy.
Lee, John D.
In two wind tunnels used for the two-dimensional airfoil tests, each wall above and below the model was modified by replacing small segments of the solid boundaries with perforated plates vented into sealed chambers. Perforated segments having approximately 40 percent open area were found to reduce the transonic wall interference to a negligible level, for a model chord-to-tunnel height ratio of 0.5. This report describes the physical arrangement and presents typical model pressure distributions to illustrate the effectiveness of the technique.
Weatherill, W. H.; Ehlers, F. E.
A finite difference method for analyzing the unsteady transonic flow about harmonically oscillating wings is discussed. The procedure is based on separating the velocity potential into steady and unsteady parts and linearizing the resulting partial differential equations for small disturbances. Initial solutions were obtained using relaxation procedures, but the solution range proved to be limited in terms of Mach number and reduced frequency. Recent two-dimensional results are presented which have been obtained with direct solution procedures in which the difference equations are solved 'all at once' and these provide reasonable correlation for practical values of Mach number and reduced frequency.
Stahara, S. S.; Spreiter, J. R.
A nonlinear, unsteady, small-disturbance theory capable of predicting inviscid transonic flows about aerodynamic configurations undergoing both rigid body and elastic oscillations was developed. The theory is based on the concept of dividing the flow into steady and unsteady components and then solving, by method of local linearization, the coupled differential equation for unsteady surface pressure distribution. The equations, valid at all frequencies, were derived for two-dimensional flows, numerical results, were obtained for two classses of airfoils and two types of oscillatory motions.
Holst, T. L.; Brown, D.
A new algorithm for generating solution-adaptive grids (SAG) about airfoil configurations embedded in transonic flow is presented. The present SAG approach uses only the airfoil surface solution to recluster grid points on the airfoil surface, i.e., the reclustering problem is one dimension smaller than the flow-field calculation problem. Special controls automatically built into the elliptic grid generation procedure are then used to obtain grids with suitable interior behavior. This concept of redistributing grid points greatly simplifies the idea of solution-adaptive grids. Numerical results indicate significant improvements in accuracy for SAG grids relative to standard grids using the same number of points.
Nakayama, A.; Chow, W. L.
A transonic flow past a boattailed afterbody under a small angle of attack was examined. It is known that the viscous effect offers significant modifications of the pressure distribution on the afterbody. Thus, the formulation for the inviscid flow was based on the consideration of a flow past a nonaxisymmetric body. The full three dimensional potential equation was solved through numerical relaxation, and quasi-axisymmetric boundary layer calculations were performed to estimate the displacement effect. It was observed again that the viscous effects were not negligible. The trend of the final results agreed well with the experimental data.
Hathaway, M. D.; Gertz, J.; Epstein, A.; Strazisar, A. J.
State of the art turbomachinery flow analysis codes are not capable of predicting the viscous flow features within turbomachinery blade wakes. Until efficient 3D viscous flow analysis codes become a reality there is therefore a need for models which can describe the generation and transport of blade wakes and the mixing process within the wake. To address the need for experimental data to support the development of such models, high response pressure measurements and laser anemometer velocity measurements were obtained in the wake of a transonic axial flow fan rotor.
Jutras, R. R.; Fost, R. B.; Chi, R. M.; Beacher, B. F.
Stall flutter is investigated by obtaining detailed quantitative steady and aerodynamic and aeromechanical measurements in a typical fan rotor. The experimental investigation is made with a 31.3 percent scale model of the Quiet Engine Program Fan C rotor system. Both subsonic/transonic (torsional mode) flutter and supersonic (flexural) flutter are investigated. Extensive steady and unsteady data on the blade deformations and aerodynamic properties surrounding the rotor are acquired while operating in both the steady and flutter modes. Analysis of this data shows that while there may be more than one traveling wave present during flutter, they are all forward traveling waves.
Santillan, S.; Selmin, V.
A methodology for solving the Euler equations in geometrically complex domains has been developed at Alenia (DVD). The general features of the Alenia multiblock system, which provides a grid generation and Euler solution capability for complex configuration, are discussed. A systematic multiblock grid generator has been used to make the grids. Numerical results compared with available experimental data in transonic vortical flow around delta wing, wing-body and wing-body-canard configurations with sharp leading edge are presented and explained in a physical context.
Zhu, Ziqiang; Xia, Zhixun; Wu, Liyi
It is known from Lighthill's exact solution of the incompressible inverse problem that in the inverse design problem, the surface pressure distribution and the free stream speed cannot both be prescribed independently. This implies the existence of a constraint on the prescribed pressure distribution. The same constraint exists at compressible speeds. Presented here is an inverse design method for transonic airfoils. In this method, the target pressure distribution contains a free parameter that is adjusted during the computation to satisfy the regularity condition. Some design results are presented in order to demonstrate the capabilities of the method.
Liang, S.-M.; Fung, K.-Y.
A numerical procedure combining the ideas of solving a modified difference equation and of adaptive mesh refinement is introduced. The numerical solution on a fixed grid is improved by using better approximations of the truncation error computed from local subdomain grid refinements. This technique is used to obtain refined solutions of steady, inviscid, transonic flow past a wedge. The effects of truncation error on the pressure distribution, wave drag, sonic line, and shock position are investigated. By comparing the pressure drag on the wedge and wave drag due to the shocks, a supersonic-to-supersonic shock originating from the wedge shoulder is confirmed.
Stanford, Bret K.; Wieseman, Carol D.; Jutte, Christine V.
Several minimum-mass optimization problems are solved to evaluate the effectiveness of a variety of novel tailoring schemes for subsonic transport wings. Aeroelastic stress and panel buckling constraints are imposed across several trimmed static maneuver loads, in addition to a transonic flutter margin constraint, captured with aerodynamic influence coefficient-based tools. Tailoring with metallic thickness variations, functionally graded materials, balanced or unbalanced composite laminates, curvilinear tow steering, and distributed trailing edge control effectors are all found to provide reductions in structural wing mass with varying degrees of success. The question as to whether this wing mass reduction will offset the increased manufacturing cost is left unresolved for each case.
Sprinkle, Danny R.; Obara, Clifford J.; Amer, Tahani R.; Leighty, Bradley D.; Carmine, Michael T.; Sealey, Bradley S.; Burkett, Cecil G.
This report describes the NASA Langley 16-Ft. Transonic Tunnel Pressure Sensitive Paint (PSP) System and presents results of a test conducted June 22-23, 2000 in the tunnel to validate the PSP system. The PSP system provides global surface pressure measurements on wind tunnel models. The system was developed and installed by PSP Team personnel of the Instrumentation Systems Development Branch and the Advanced Measurement and Diagnostics Branch. A discussion of the results of the validation test follows a description of the system and a description of the test.
Phillips, Pamela S.; Waggoner, Edgar G.; Campbell, Richard L.
An automated transonic design code has been developed which modifies an initial airfoil or wing in order to generate a specified pressure distribution. The design method uses an iterative approach that alternates between a potential-flow analysis and a design algorithm that relates changes in surface pressure to changes in geometry. The analysis code solves an extended small-disturbance potential-flow equation and can model a fuselage, pylons, nacelles, and a winglet in addition to the wing. A two-dimensional option is available for airfoil analysis and design. Several two- and three-dimensional test cases illustrate the capabilities of the design code.
Campbell, Richard L.; Smith, Leigh A.
The present method for the design of transonic airfoils and wings employs a predictor/corrector approach in which an analysis code calculates the flowfield for an initial geometry, then modifies it on the basis of the difference between calculated and target pressures. This allows the design method to be straightforwardly coupled with any existing analysis code, as presently undertaken with several two- and three-dimensional potential flow codes. The results obtained indicate that the method is robust and accurate, even in the cases of airfoils with strongly supercritical flow and shocks. The design codes are noted to require computational resources typical of current pure-inverse methods.
Stahara, S. S.
This code solves the two-dimensional, transonic, small-disturbance equations for flow past lifting airfoils in both free air and various wind-tunnel environments by using a variant of the finite-difference method. A description of the theoretical and numerical basis of the code is provided, together with complete operating instructions and sample cases for the general user. In addition, a programmer's manual is also presented to assist the user interested in modifying the code. Included in the programmer's manual are a dictionary of subroutine variables in common and a detailed description of each subroutine.
Liou, William W.; Shih, Tsan-Hsing
Solutions of the Favre-averaged Navier-Stokes equations for two well-documented transonic turbulent flows are compared in detail with existing experimental data. While the boundary layer in the first case remains attached, a region of extensive flow separation has been observed in the second case. Two recently developed k-epsilon, two-equation, eddy-viscosity models are used to model the turbulence field. These models satisfy the realizability constraints of the Reynolds stresses. Comparisons with the measurements are made for the wall pressure distribution, the mean streamwise velocity profiles, and turbulent quantities. Reasonably good agreement is obtained with the experimental data.
Kittleson, John K.; Yu, Yung H.
Holographic interferometry is used to record interferograms of the flow near a hovering transonic rotor blade. A pulsed ruby laser recorded 40 interferograms with a 2 ft dia. view field near the model rotor blade tip operating at a tip Mach number of 0.90. The experimental procedure is presented and example interferograms recorded in the rotor's tip path plane. In addition, a method currently being pursued to obtain quantitative flow information using computer assisted tomography (CAT) with the holographic interferogram data, is outlined.
Couch, L. M.
Efforts to develop a near sonic transport have placed renewed emphasis on obtaining accurate aerodynamic force and pressure data in the near sonic speed range. Comparison of wind-tunnel and flight data obtained for a blunt-nose body of revolution showed significant discrepancies in drag levels near Mach 1 - apparently due to wind-tunnel wall interference. Subsequent tests of geometrically similar bodies of revolution showed that increasing the model-to-test-section blockage ratio from 0.00017 to 0.0043 resulted in altered drag curve shapes, delayed drag divergence, and 'transonic creep' from subsonic drag levels due to increased wall interference.
Doggett, Robert V.
The Transonic Dynamics Tunnel (TDT) at the National Aeronautics and Space Administration's (NASA) Langley Research Center began research operations in early 1960. Since that time, over 600 tests have been conducted, primarily in the discipline of aeroelasticity. This paper presents a bibliography of the publications that contain data from these tests along with other reports that describe the facility, its capabilities, testing techniques, and associated research equipment. The bibliography is divided by subject matter into a number of categories. An index by author's last name is provided.
Miranda, Luis R.
The application of computational fluid dynamics (CFD) to fighter aircraft design and development is discussed. Methodology requirements for the aerodynamic design of fighter aircraft are briefly reviewed. The state-of-the-art of computational methods for transonic flows in the light of these requirements is assessed and the techniques found most adequate for the subject application are identified. Highlights from some proof-of-feasibility Euler and Navier-Stokes computations about a complete fighter aircraft configuration are presented. Finally, critical issues and opportunities for design application of CFD are discussed.
past an NACA 64A010 airfoil in a slotted wall wind- tunnel. The drag rise due to shock wave formation (CD0 friction drag) is fairly well represented...Some results of calculations are shown in Fig. 10 and 11 for a parabolic arc airfoil 9 and a NACA 001210. For the first the sonic line and shock...Number na- Shock Attachment NACA TN 3225, 1954. ..................." a S. . . . -19- Figure Titles Fig. I Transonic Flow past an Airfoil Fig. 2 Three
Mann, M. J.; Mercer, C. E.
A transonic computational analysis method and a transonic design procedure have been used to design the wing and the canard of a forward-swept-wing fighter configuration for good transonic maneuver performance. A model of this configuration was tested in the Langley 16-Foot Transonic Tunnel. Oil-flow photographs were obtained to examine the wind flow patterns at Mach numbers from 0.60 to 0.90. The transonic theory gave a reasonably good estimate of the wing pressure distributions at transonic maneuver conditions. Comparison of the forward-swept-wing configuration with an equivalent aft-swept-wing-configuration showed that, at a Mach number of 0.90 and a lift coefficient of 0.9, the two configurations have the same trimmed drag. The forward-swept wing configuration was also found to have trimmed drag levels at transonic maneuver conditions which are comparable to those of the HiMAT (highly maneuverable aircraft technology) configuration and the X-29 forward-swept-wing research configuration. The configuration of this study was also tested with a forebody strake.
Chan, David T.; Hooker, John R.; Wick, Andrew; Plumley, Ryan W.; Zeune, Cale H.; Ol, Michael V.; DeMoss, Joshua A.
A wind tunnel investigation of a 0.04-scale model of the Lockheed Martin Hybrid Wing Body (HWB) with Over Wing Nacelles (OWN) air mobility transport configuration was conducted in the National Transonic Facility at the NASA Langley Research Center under a collaborative partnership between NASA, the Air Force Research Laboratory, and Lockheed Martin Aeronautics Company. The wind tunnel test sought to validate the transonic aerodynamic performance of the HWB and to validate the efficiency benefits of the OWN installation as compared to the traditional under-wing installation. The semispan HWB model was tested in a clean wing configuration and also tested with two different nacelles representative of a modern turbofan engine and a future advanced high bypass ratio engine. The nacelles were installed in three different locations with two over-wing positions and one under-wing position. Five-component force and moment data, surface static pressure data, and aeroelastic deformation data were acquired. For the cruise configuration, the model was tested in an angle-of-attack range between -2 and 10 degrees at free-stream Mach numbers from 0.3 to 0.9 and at unit Reynolds numbers between 8 and 39 million per foot, achieving a maximum of 80% of flight Reynolds numbers across the Mach number range. The test results validated pretest computational fluid dynamic (CFD) simulations of the HWB performance including the OWN benefit and the results also exhibited excellent transonic drag data repeatability to within +/-1 drag count. This paper details the experimental setup and model overview, presents some sample data results, and describes the facility improvements that led to the success of the test.
Luckring, J. M.
A 65 deg delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at transonic speeds (M = 0.85) from this data set. The results show significant effects of both these parameters on the onset and progression of leading- edge vortex separation.
Luckring, J. M.
A 65 deg delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at transonic speeds (M=0.85) from this data set. The results show significant effects of both these parameters on the onset and progression of leading-edge vortex separation.
Luckring, J. M.
A 65 degree delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at transonic speeds (M = 0.85) from this data set. The results show significant effects of both these parameters on the onset and progression of leading edge vortex separation.
Erickson, L. L.; Strande, S. M.
The surface integral terms in Green's third identity are often used to solve the Prandtl-Glauert (linear potential-flow) equation with panel methods. This can be done, as in the PAN AIR code, for either subsonic or supersonic flow about complete aircraft. The extension to transonic flow is suggested by the volume integral terms of Green's third identity. The mathematical basis for this extension, without the use of body-fitted grids, is presented. Supercritical transonic results computed from a two-dimensional transonic PAN AIR research code demonstrate the method.
Abramova, K. A.; Petrov, A. V.; Potapchick, A. V.; Soudakov, V. G.
Numerical and experimental investigations of transonic buffet control by tangential jet blowing are presented. To suppress the shock-induced boundary layer separation and the buffet at transonic speeds, compressed air jet is blown through a small slot nozzle tangentially to the upper surface of the supercritical airfoil. Numerical simulations were carried out on the basis of the unsteady Reynolds averaged Navier-Stokes (URANS) equations. Experimental studies of the tangential jet blowing were performed in the transonic wind tunnel T-112 of TsAGI. Results show that the jet moves the shock downstream, increases lift, suppresses flow separation under shock foot and delays buffet onset.
Newman, P. A.; Anderson, E. C.
An analytical procedure is discussed for designing wall shapes for streamlined, nonporous, two-dimensional, transonic wind tunnels. It is based upon currently available 2-D inviscid transonic and boundary layer analysis computer programs. Predicted wall shapes are compared with experimental data obtained from the NASA Langley 6 by 19 inch Transonic Tunnel where the slotted walls were replaced by flexible nonporous walls. Comparisons are presented for the empty tunnel operating at a Mach number of 0.9 and for a supercritical test of an NACA 0012 airfoil at zero lift. Satisfactory agreement is obtained between the analytically and experimentally determined wall shapes.
Paryz, Roman W.
Several projects have been completed or are nearing completion at the NASA Langley Research Center (LaRC) National Transonic Facility (NTF). The addition of a Model Flow-Control/Propulsion Simulation test capability to the NTF provides a unique, transonic, high-Reynolds number test capability that is well suited for research in propulsion airframe integration studies, circulation control high-lift concepts, powered lift, and cruise separation flow control. A 1992 vintage Facility Automation System (FAS) that performs the control functions for tunnel pressure, temperature, Mach number, model position, safety interlock and supervisory controls was replaced using current, commercially available components. This FAS upgrade also involved a design study for the replacement of the facility Mach measurement system and the development of a software-based simulation model of NTF processes and control systems. The FAS upgrades were validated by a post upgrade verification wind tunnel test. The data acquisition system (DAS) upgrade project involves the design, purchase, build, integration, installation and verification of a new DAS by replacing several early 1990's vintage computer systems with state of the art hardware/software. This paper provides an update on the progress made in these efforts. See reference 1.
Stern, Eric; Barnhardt, Michael; Venkatapathy, Ethiraj; Candler, Graham; Prabhu, Dinesh
A numerical investigation of transonic flow around a mechanically deployable entry system being considered for a robotic mission to Venus has been performed, and preliminary results are reported. The flow around a conceptual representation of the vehicle geometry was simulated at discrete points along a ballistic trajectory using Detached Eddy Simulation (DES). The trajectory points selected span the low supersonic to transonic regimes with freestream Mach numbers from 1:5 to 0:8, and freestream Reynolds numbers (based on diameter) between 2:09 x 10(exp 6) and 2:93 x 10(exp 6). Additionally, the Mach 0:8 case was simulated at angles of attack between 0 and 5 . Static aerodynamic coefficients obtained from the data show qualitative agreement with data from 70deg sphere-cone wind tunnel tests performed for the Viking program. Finally, the effect of choices of models and numerical algorithms is addressed by comparing the DES results to those using a Reynolds Averaged Navier-Stokes (RANS) model, as well as to results using a more dissipative numerical scheme.
Johnson, D. A.; Bachalo, W. D.; Moddaress, D.
As a further demonstration of the capabilities of laser velocity in compressible aerodynamics, measurements obtained in a Mach 2.9 separated turbulent boundary layer and in the transonic flow past a two-dimensional airfoil section are presented and compared to data realized by conventional techniques. In the separated-flow study, the comparisons were made against pitot-static pressure data. Agreement in mean velocities was realized where the pressure measurements could be considered reliable; however, in regions of instantaneous reverse velocities, the laser results were found to be consistent with the physics of the flow whereas the pressure data were not. The laser data obtained in regions of extremely high turbulence suggest that velocity biasing does not occur if the particle occurrence rate is low relative to the turbulent fluctuation rate. Streamwise turbulence intensities are also presented. In the transonic airfoil study, velocity measurements obtained immediately outside the upper surface boundary layer of a 6-inch chord MACA 64A010 airfoil are compared to edge velocities inferred from surface pressure measurements. For free-stream Mach numbers of 0.6 and 0.8, the agreement in results was very good. Dual scatter optical arrangements in conjunction with a single particle, counter-type signal processor were employed in these investigations. Half-micron-diameter polystyrene spheres and naturally occurring condensed oil vapor acted as light scatterers in the two respective flows. Bragg-cell frequency shifting was utilized in the separated flow study.
Tracy, Maureen B.; Plentovich, E. B.
An experimental investigation was conducted in the Langley 8-Foot Transonic Pressure Tunnel to determine the flow characteristics of rectangular cavities with varying relative dimensions at subsonic and transonic speeds. Cavities were tested with width-to-depth ratios of 1, 4, 8, and 16 for length-to-depth ratios l/h of 1 through 17.5. The maximum cavity depth was 2.4 in., and the turbulent boundary layer approaching the cavity was approximately 0.5 in. thick. Unsteady- and mean static-pressure measurements were made at free-stream Mach numbers from 0.20 to 0.95 at a unit Reynolds number per foot of approximately 3 x 10(exp 6); however, only unsteady-pressure results are presented in this paper. Results indicate that as l/h increases, cavity flows changed from resonant to nonresonant with resonant amplitudes decreasing gradually. Resonant spectra are obtained largely in cavities with mean static-pressure distributions characteristic of open and transitional flows. Resonance sometimes occurred for closed flow. Increasing cavity width or decreasing cavity depth while holding l/h fixed had the effect of increasing resonant amplitudes and sometimes induced resonance. The effects due to changes in width are more pronounced. Decreasing Mach number has the effect of broadening the resonances.
Mavriplis, D. J.; Levy, David W.
This paper summarizes the results obtained with the NSU-3D unstructured multigrid solver for the AIAA Drag Prediction Workshop held in Anaheim, CA, June 2001. The test case for the workshop consists of a wing-body configuration at transonic flow conditions. Flow analyses for a complete test matrix of lift coefficient values and Mach numbers at a constant Reynolds number are performed, thus producing a set of drag polars and drag rise curves which are compared with experimental data. Results were obtained independently by both authors using an identical baseline grid and different refined grids. Most cases were run in parallel on commodity cluster-type machines while the largest cases were run on an SGI Origin machine using 128 processors. The objective of this paper is to study the accuracy of the subject unstructured grid solver for predicting drag in the transonic cruise regime, to assess the efficiency of the method in terms of convergence, cpu time, and memory, and to determine the effects of grid resolution on this predictive ability and its computational efficiency. A good predictive ability is demonstrated over a wide range of conditions, although accuracy was found to degrade for cases at higher Mach numbers and lift values where increasing amounts of flow separation occur. The ability to rapidly compute large numbers of cases at varying flow conditions using an unstructured solver on inexpensive clusters of commodity computers is also demonstrated.
Cosentino, G. B.; Holst, T. L.
A computationally efficient and versatile technique for use in the design of advanced transonic wing configurations has been developed. A reliable and fast transonic wing flow-field analysis program, TWING, has been coupled with a modified quasi-Newton method, unconstrained optimization algorithm, QNMDIF, to create a new design tool. Fully three-dimensional wing designs utilizing both specified wing pressure disributions and drag-to-lift ratio minimization as design objectives are demonstrated. Because of the high computational efficiency of each of the components of the design code, in particular the vectorization of TWING and the high speed of the Cray X-MP vector computer, the computer time required for a typical wing design is reduced by approximately an order of magnitude over previous methods. In the results presented here, this computed wave drag has been used as the quantity to be optimized (minimized) with great success, yielding wing designs with nearly shock-free (zero wave drag) pressure distributions and very reasonable wing section shapes.
Cosentino, G. B.; Holst, T. L.
A computationally efficient and versatile technique for use in the design of advanced transonic wing configurations has been developed. A reliable and fast transonic wing flow-field analysis program, TWING, has been coupled with a modified quasi-Newton method, unconstrained optimization algorithm, QNMDIF, to create a new design tool. Fully three-dimensional wing designs utilizing both specified wing pressure distributions and drag-to-lift ration minimization as design objectives are demonstrated. Because of the high computational efficiency of each of the components of the design code, in particular the vectorization of TWING and the high speed of the Cray X-MP vector computer, the computer time required for a typical wing design is reduced by approximately an order of magnitude over previous methods. In the results presented here, this computed wave drag has been used as the quantity to be optimized (minimized) with great success, yielding wing designs with nearly shock-free (zero wave drag) pressure distributions and very reasonable wing section shapes.
Bendiksen, Oddvar O.
This paper presents an overview of recent applications of Eulerian-Lagrangian computational schemes in simulating transonic flutter instabilities. This approach, the fluid-structure system is treated as a single continuum dynamics problem, by switching from an Eulerian to a Lagrangian formulation at the fluid-structure boundary. This computational approach effectively eliminates the phase integration errors associated with previous methods, where the fluid and structure are integrated sequentially using different schemes. The formulation is based on Hamilton's Principle in mixed coordinates, and both finite volume and finite element discretization schemes are considered. Results from numerical simulations of transonic flutter instabilities are presented for isolated wings, thin panels, and turbomachinery blades. The results suggest that the method is capable of reproducing the energy exchange between the fluid and the structure with significantly less error than existing methods. Localized flutter modes and panel flutter modes involving traveling waves can also be simulated effectively with no a priori knowledge of the type of instability involved.
Lyrintzis, A. S.; Xue, Y.
Transonic Blade-Vortex Interactions (BVI) are simulated numerically and the noise mechanisms are investigated. The 2-D high frequency transonic small disturbance equation is solved numerically (VTRAN2 code). An Alternating Direction Implicit (ADI) scheme with monotone switches is used; viscous effects are included on the boundary and the vortex is simulated by the cloud-in-cell method. The Kirchoff method is used for the extension of the numerical 2-D near field aerodynamic results to the linear acoustic 3-D far field. The viscous effect (shock/boundary layer interaction) on BVI is investigated. The different types of shock motion are identified and compared. Two important disturbances with different directivity exist in the pressure signal and are believed to be related to the fluctuating lift and drag forces. Noise directivity for different cases is shown. The maximum radiation occurs at an angle between 60 and 90 deg below the horizontal for an airfoil fixed coordinate system and depends on the details of the airfoil shape. Different airfoil shapes are studied and classified according to the BVI noise produced.
Nicke, E.; Steinert, W.; Weber, A.; Starken, H.
Starting from two existing cascades a new one was developed for the same inlet and exit flow conditions but increased pitch to chord ratio: inlet Mach number M(sub 1) = 1.09, flow turning Theta = 14 degrees, and pitch to chord ratio s/c = 0.8. Perspective research goal is the determination of the loading limits of transonic compressor cascades. The design was performed in the direct mode using a two dimensional Euler code coupled iteratively with an integral boundary layer code generating a profile Mach number distribution that avoids boundary layer separation. The experiments in a transonic cascade wind tunnel revealed performance, profile, Mach number distribution, and wake traverse data. A laser, two focus measurement technique was used to analyze the flow field upstream of the cascade and around the leading edge. The experimental results are analyzed by the Euler code. To improve the reliability of experimental as well as numerical inlet flow angle determination, the influence of conventional inlet boundary conditions versus so called nonreflecting boundary conditions was tested and verified using the L2F results.
Wood, Jerry R.; Strazisar, Anthony J.; Simonyi, P. Susan
Shock structure measurements acquired in a low aspect ratio transonic fan rotor are presented and analyzed. The rotor aspect ratio is 1.56 and the design tip relative Mach number is 1.38. The rotor flowfield was surveyed at near maximum efficiency and near stall operating conditions. Intra-blade velocity measurements acquired with a laser fringe anemometer on blade-to-blade planes in the supersonic region from 10 to 60 percent span are presented. The three-dimensional shock surface determined from the velocity measurments is used to determine the shock surface normal Mach number in order to properly calculate the ideal shock jump conditions. The ideal jump conditions are calculated based upon the Mach numbers measured on a surface of revolution and based upon the normal Mach number to indicate the importance of accounting for shock three dimensionality in turbomachinery design. Comparison of the shock locations with those predicted by a 3-D Euler code showed very good agreement and indicated the usefulness of integrating computational and experimental work to enhance the understanding of the flow physics occurring in transonic turbomachinery passages.
Chima, Rodrick V.
A 3-D flow analysis code was used to compute the design speed operating line of a transonic fan rotor, and the results were compared with experimental data. The code is an explicit finite difference code with an algebraic turbulence model. The transonic fan, called rotor 67, was tested experimentally at NASA-Lewis with conventional aerodynamic probes and with user anemometry and was included as one of the AGARD test cases for the computation of internal flows. The experimental data are described. Maps of total pressure ratio and adiabatic efficiency versus mass flow were computed and are compared with the experimental maps, with good agreement. Detailed comparisons between calculations and experiment are made at two operating points, one near peak efficiency and the other near stall. Blade-to-blade contour plots are used to show the shock structure. Comparisons of circumferentially integrated flow quantities downstream of the rotor show spanwise distributions of several aerodynamic parameters. Calculated Mach number distributions are compared with laser anemometer data within the blade row and the wake to quantify the accuracy of the calculations. Particle traces are used to show the nature of secondary flow.
Chima, Rodrick V.
A 3-D flow analysis code was used to compute the design speed operating line of a transonic fan rotor, and the results were compared with experimental data. The code is an explicit finite difference code with an algebraic turbulence model. The transonic fan, called Rotor 67, was tested experimentally at NASA Lewis conventional aerodynamic probes and with user anemometry and was included as one of the AGARD test cases for the computation of internal flows. The experimental data are described. Maps of total pressure ratio and adiabatic efficiency vs mass flow were computed and are compared with the experimental maps, with good agreement. Detailed comparisons between calculations and experiment are made at two operating points, one near peak efficiency and the other near stall. Blade-to-blade contour plots are used to show the shock structure. Comparisons of circumferentially integrated flow quantities downstream of the rotor show spanwise distributions of several aerodynamic parameters. Calculated Mach number distributions are compared with laser anemometer data within the blade row and the wake to quantify the accuracy of the calculations. Particle traces are used to show the nature of secondary flow.
King, Rudolph A.; Andino, Marlyn Y.; Melton, Latunia; Eppink, Jenna; Kegerise, Michael A.
Recent flow measurements have been acquired in the National Transonic Facility to assess the test-section unsteady flow environment. The primary purpose of the test is to determine the feasibility of the facility to conduct laminar-flow-control testing and boundary-layer transition-sensitive testing at flight-relevant operating conditions throughout the transonic Mach number range. The facility can operate in two modes, warm and cryogenic test conditions for testing full and semispan-scaled models. Data were acquired for Mach and unit Reynolds numbers ranging from 0.2 less than or equal to M less than or equal to 0.95 and 3.3 × 10(exp 6) less than Re/m less than 220×10(exp 6) collectively at air and cryogenic conditions. Measurements were made in the test section using a survey rake that was populated with 19 probes. Roll polar data at selected conditions were obtained to look at the uniformity of the flow disturbance field in the test section. Data acquired included mean total temperatures, mean and fluctuating static/total pressures, and mean and fluctuating hot-wire measurements. This paper focuses primarily on the unsteady pressure and hot-wire results. Based on the current measurements and previous data, an assessment was made that the facility may be a suitable facility for ground-based demonstrations of laminar-flow technologies at flight-relevant conditions in the cryogenic mode.
King, Rudolph A.; Andino, Marlyn Y.; Melton, Latunia; Eppink, Jenna; Kegerise, Michael A.; Tsoi, Andrew
Recent flow measurements have been acquired in the National Transonic Facility (NTF) to assess the unsteady flow environment in the test section. The primary purpose of the test is to determine the feasibility of the NTF to conduct laminar-flow-control testing and boundary-layer transition sensitive testing. The NTF can operate in two modes, warm (air) and cold/cryogenic (nitrogen) test conditions for testing full and semispan scaled models. The warm-air mode enables low to moderately high Reynolds numbers through the use of high tunnel pressure, and the nitrogen mode enables high Reynolds numbers up to flight conditions, depending on aircraft type and size, utilizing high tunnel pressure and cryogenic temperatures. NASA's Environmentally Responsible Aviation (ERA) project is interested in demonstrating different laminar-flow technologies at flight-relevant operating conditions throughout the transonic Mach number range and the NTF is well suited for the initial ground-based demonstrations. Roll polar data at selected test conditions were obtained to look at the uniformity of the flow disturbance field in the test section. Data acquired from the rake probes included mean total temperatures, mean and fluctuating static/total pressures, and mean and fluctuating hot-wire measurements. . Based on the current measurements and previous data, an assessment was made that the NTF is a suitable facility for ground-based demonstrations of laminar-flow technologies at flight-relevant conditions in the cryogenic mode.
Piatak, David J.; Sekula, Martin K.; Rausch, Russ D.
It is necessary to define the launch vehicle buffet loads to ensure that structural components and vehicle subsystems possess adequate strength, stress, and fatigue margins when the vehicle structural dynamic response to buffet forcing functions are considered. In order to obtain these forcing functions, the accepted method is to perform wind-tunnel testing of a rigid model instrumented with hundreds of unsteady pressure transducers designed to measure the buffet environment across the desired frequency range. The buffet wind-tunnel test program for the Ares Crew Launch Vehicle employed 3.5 percent scale rigid models of the Ares I and Ares I-X launch vehicles instrumented with 256 unsteady pressure transducers each. These models were tested at transonic conditions at the Transonic Dynamics Tunnel at NASA Langley Research Center. The ultimate deliverable of the Ares buffet test program are buffet forcing functions (BFFs) derived from integrating the measured fluctuating pressures on the rigid wind-tunnel models. These BFFs are then used as input to a multi-mode structural analysis to determine the vehicle response to buffet and the resulting buffet loads and accelerations. This paper discusses the development of the Ares I and I-X rigid buffet model test programs from the standpoint of model design, instrumentation system design, test implementation, data analysis techniques to yield final products, and presents normalized sectional buffet forcing function root-mean-squared levels.
Chang, I C
With the aim of developing a fast and accurate computer code for predicting the aerodynamic forces needed for a flutter analysis, some basic concepts in computational transonics are reviewed. The unsteady transonic flow past airfoils in rigid body motion is adequately described by the potential flow equation as long as the boundary layer remains attached. The two dimensional unsteady transonic potential flow equation in quasilinear form with first order radiation boundary conditions is solved by an alternating direction implicit scheme in an airfoil attached sheared parabolic coordinate system. Numerical experiments show that the scheme is very stable and is able to resolve the higher nonlinear transonic effects for filter analysis within the context of an inviscid theory.
Holst, T. L.
Numerical solution techniques for solving transonic flow fields governed by the full potential equation are discussed. In a general sense relaxation schemes suitable for the numerical solution of elliptic partial differential equations are presented and discussed with emphasis on transonic flow applications. The presentation can be divided into two general categories: An introductory treatment of the basic concepts associated with the numerical solution of elliptic partial differential equations and a more advanced treatment of current procedures used to solve the full potential equation for transonic flow fields. The introductory material is presented for completeness and includes a brief introduction (Chapter 1), governing equations (Chapter 2), classical relaxation schemes (Chapter 3), and early concepts regarding transonic full potential equation algorithms (Chapter 4).
Batina, John T.
Modifications to unsteady transonic small disturbance theory to include entropy and vorticity effects are presented. The modifications were implemented in the CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code. The code permits the aeroelastic analysis of complete aircraft configurations in the flutter critical transonic speed range. Entropy and vorticity effects were incorporated within the solution procedure to more accurately analyze flows with strong shock waves. The modified code includes these effects while retaining the relative simplicity and cost efficiency of the TSD formulation. Detailed descriptions are presented of the entropy and vorticity modifications along with calculated results and comparisons which assess the modified theory. These results are in good agreement with parallel Euler calculations and with experimental data. Therefore, the present method now provides the aeroelastician with an affordable capability to analyze relatively difficult transonic flows without having to solve the computationally more expensive Euler equations.
documented that uses commercial-off-the-shelf software (MATLAB, SolidWorks , and ANSYS-CFX) for the geometric rendering and analysis of a transonic axial...that uses commercial-off-the- shelf software (MATLAB, SolidWorks , and ANSYS-CFX) for the geometric rendering and analysis of a transonic axial...Postgraduate School NPSMF NPS Military Fan NTG No Tip Gap PS Pressure Side SB Splitter Blade SS Suction Side SolidWorks A solid modeling
Van Leer, Bram; Lee, Wen-Tzong; Powell, Kenneth G.
A prototype scheme that produces perfectly smooth transonic solutions to nozzle-flow problems is derived and tested. The basic upwind scheme is described as well as satisfying the entropy condition, treatment of the source term, and numerical verification. The analysis yielded a numerical flux function for use near a sonic point, which is based on a full model of a transonic expansion wave, and a matched treatment for the source term.
Couston, M.; Angelini, J. J.; Mulak, P.
Transonic nonsteady flows are of large practical interest. Aeroelastic instability prediction, control figured vehicle techniques or rotary wings in forward flight are some examples justifying the effort undertaken to improve knowledge of these problems is described. The numerical solution of these problems under the potential flow hypothesis is described. The use of an alternating direction implicit scheme allows the efficient resolution of the two dimensional transonic small perturbations equation.
AFRL-RQ-WP-TP-2014-0168 SERPENTINE DIFFUSER PERFORMANCE WITH EMPHASIS ON FUTURE INTRODUCTION TO A TRANSONIC FAN (POSTPRINT) Chase A...June 2012 – 01 December 2012 4. TITLE AND SUBTITLE SERPENTINE DIFFUSER PERFORMANCE WITH EMPHASIS ON FUTURE INTRODUCTION TO A TRANSONIC FAN...subject to copyright protection in the United States. 14. ABSTRACT A serpentine diffuser for embedded propulsion systems was characterized in terms
Carlson, Leland A.
The primary accomplishments of the project are as follows: (1) Using the transonic small perturbation equation as a flowfield model, the project demonstrated that the quasi-analytical method could be used to obtain aerodynamic sensitivity coefficients for airfoils at subsonic, transonic, and supersonic conditions for design variables such as Mach number, airfoil thickness, maximum camber, angle of attack, and location of maximum camber. It was established that the quasi-analytical approach was an accurate method for obtaining aerodynamic sensitivity derivatives for airfoils at transonic conditions and usually more efficient than the finite difference approach. (2) The usage of symbolic manipulation software to determine the appropriate expressions and computer coding associated with the quasi-analytical method for sensitivity derivatives was investigated. Using the three dimensional fully conservative full potential flowfield model, it was determined that symbolic manipulation along with a chain rule approach was extremely useful in developing a combined flowfield and quasi-analytical sensitivity derivative code capable of considering a large number of realistic design variables. (3) Using the three dimensional fully conservative full potential flowfield model, the quasi-analytical method was applied to swept wings (i.e. three dimensional) at transonic flow conditions. (4) The incremental iterative technique has been applied to the three dimensional transonic nonlinear small perturbation flowfield formulation, an equivalent plate deflection model, and the associated aerodynamic and structural discipline sensitivity equations; and coupled aeroelastic results for an aspect ratio three wing in transonic flow have been obtained.
weak solutions. The linear theory is deficient in predicting important features of transonic flow outside airfoils in low reduced frequency motion...terms. The term t is substituted n+l adn -i n n+l +n-i by the mean of and , i.e., = 2 = -(u 2 +2uv y+V 2 $ q2 xx xyq 1 2uF v2-2-5 n+l n-I(v2 + 2uvT + 2 ( I...approximate for thie advection equation. Our approximate factorizatio. says that (2) can .) factored as (1+tuQ 14tv )*N - -. tu )(l-.ltvD ) M - 2.’,t (uP DvD Ki
Yetter, J. A.; Evelyn, G. B.; Mercer, C.
The design of the propulsion system installation affects strongly the total drag and overall performance of an aircraft, and the concept, placement, and integration details of the exhaust nozzle are major considerations in the configuration definition. As part of the NASA Supersonic Cruise Research (SCR) program, a wind tunnel test program has been conducted to investigate exhaust nozzle-airframe interactions at transonic speeds. First phase testing is to establish guidelines for follow-on testing. A summary is provided of the results of first phase testing, taking into account the test approach, the effect of nozzle closure on aircraft aerodynamic characteristics, nozzle installation effects and nacelle interference drag, and an analytical study of the effects of nozzle closure on the aircraft.
Burner, A. W.; Snow, W. L.; Goad, W. K.
A photogrammetric closed circuit television system to measure model deformation at the National Transonic Facility is described. The photogrammetric approach was chosen because of its inherent rapid data recording of the entire object field. Video cameras are used to acquire data instead of film cameras due to the inaccessibility of cameras which must be housed within the cryogenic, high pressure plenum of this facility. A rudimentary theory section is followed by a description of the video-based system and control measures required to protect cameras from the hostile environment. Preliminary results obtained with the same camera placement as planned for NTF are presented and plans for facility testing with a specially designed test wing are discussed.
Jameson, A.; Caughey, D. A.
A numerical method is presented for analyzing the transonic potential flow past a lifting, swept wing. A finite difference approximation to the full potential equation is solved in a coordinate system which is nearly conformally mapped from the physical space in planes parallel to the symmetry plane, and reduces the wing surface to a portion of one boundary of the computational grid. A coordinate invariant, rotated difference scheme is used, and the difference equations are solved by relaxation. The method is capable of treating wings of arbitrary planform and dihedral, although approximations in treating the tips and vortex sheet make its accuracy suspect for wings of small aspect ratio. Comparisons of calculated results with experimental data are shown for examples of both conventional and supercritical transport wings. Agreement is good for both types, but it was found necessary to account for the displacement effect of the boundary layer for the supercritical wing, presumably because of its greater sensitivity to changes in effective geometry.
Transonic sting interference has been studied in a supersonic wind tunnel to obtain free flight and sting support data on identical models. The two principal configurations, representing fuselage bodies, were cigar shaped with tail fins. The others were a sharp 10-deg cone, a sphere, and a blunt entry body. Comparative data indicated that the sting had an appreciable effect on drag for the fuselage-like configurations; drag rise occurred 0.02 Mach number earlier in free flight, and drag level was 15% greater. The spheres and the blunt bodies were insensitive to the presence of stings regardless of their size. The 10-deg cones were in between, experiencing no drag difference with a minimum diameter sting, but a moderate difference with the largest diameter sting tested. All data tend to confirm the notion that for the more slender bodies the sting not only affects flow but the forebody flow as well.
Dizene, R; Charbonnier, J M; Dorignac, E; Lablanc, R
An extensive study devoted to modelling blade cooling was undertaken at CEAT a few years ago in collaboration with SNECMA. For the turbomachinery applications, an experimental configuration of a turbulent boundary layer with heat transfer was studied for compressible and incompressible flows. The research presented here is a part of that study and this paper reports on the experimental results of an investigation concerned with a row of transonic jets interacting with a transverse flow. In many applications, the cooling layer does not emerge onto the surface from a tangential slot but comes from a slot normal to or inclined to what is otherwise a flush surface. In this case the freestream interacts with the coolant flow. The secondary (jet) flow is introduced at an angle of 45 degrees to the mainstream flow direction. Visualization studies using the surface flow patterns and surface temperature flow patterns are reported and discussed.
Vatsa, V. N.
An explicit multistage Runge-Kutta type time-stepping scheme is used for solving the three-dimensional, compressible, thin-layer Navier-Stokes equations. A finite-volume formulation is employed to facilitate treatment of complex grid topologies encountered in three-dimensional calculations. Convergence to steady state is expedited through usage of acceleration techniques. Further numerical efficiency is achieved through vectorization of the computer code. The accuracy of the overall scheme is evaluated by comparing the computed solutions with the experimental data for a finite wing under different test conditions in the transonic regime. A grid refinement study ir conducted to estimate the grid requirements for adequate resolution of salient features of such flows.
Sanz, J. M.
A boundary value problem for the Tricomi equation was studied in connection with transonic gas dynamics. The transformed equation delta u plus 1/3Y u sub Y equals 0 in canonical coordinates was considered in the complex domain of two independent complex variables. A boundary value problem was then set by prescribing the real part of the solution on the boundary of the real unit circle. The Dirichlet problem in the upper unit semicircle with vanishing values of the solution at Y = 0 was solved explicitly in terms of the hypergeometric function for the more general Euler-Poisson-Darboux equation. An explicit representation of the solution was also given for a mixed Dirichlet and Neumann problem for the same equation and domain.
Yetter, J. A.; Salemann, V.; Sussman, M. B.
A wind tunnel investigation was conducted to determine the local inlet flow field characteristics of an advanced tactical supersonic cruise airplane. A data base for the development and validation of analytical codes directed at the analysis of inlet flow fields for advanced supersonic airplanes was established. Testing was conducted at the NASA-Langley 16-foot Transonic Tunnel at freestream Mach numbers of 0.6 to 1.20 and angles of attack from 0.0 to 10.0 degrees. Inlet flow field surveys were made at locations representative of wing (upper and lower surface) and forebody mounted inlet concepts. Results are presented in the form of local inlet flow field angle of attack, sideflow angle, and Mach number contours. Wing surface pressure distributions supplement the flow field data.
Popernack, Thomas G., Jr.; Sydnor, George H.
Productivity gains have recently been made at the National Transonic Facility wind tunnel at NASA Langley Research Center. A team was assigned to assess and set productivity goals to achieve the desired operating cost and output of the facility. Simulations have been developed to show the sensitivity of selected process productivity improvements in critical areas to reduce overall test cycle times. The improvements consist of an expanded liquid nitrogen storage system, a new fan drive, a new tunnel vent stack heater, replacement of programmable logic controllers, an increased data communications speed, automated test sequencing, and a faster model changeout system. Where possible, quantifiable results of these improvements are presented. Results show that in most cases, improvements meet the productivity gains predicted by the simulations.
Ball, C. L.; Steinke, R. J.; Newman, F. A.
The development of the transonic multistage compressor is reviewed. Changing trends in design and performance parameters are noted. These changes are related to advances in compressor aerodynamics, computational fluid mechanics and other enabling technologies. The parameters normally given to the designer and those that need to be established during the design process are identified. Criteria and procedures used in the selection of these parameters are presented. The selection of tip speed, aerodynamic loading, flowpath geometry, incidence and deviation angles, blade/vane geometry, blade/vane solidity, stage reaction, aerodynamic blockage, inlet flow per unit annulus area, stage/overall velocity ratio, and aerodynamic losses are considered. Trends in these parameters both spanwise and axially through the machine are highlighted. The effects of flow mixing and methods for accounting for the mixing in the design process are discussed.
Underwood, Pamela J.; Owens, Lewis R.; Wahls, Richard A.; Williams, Susan
This paper describes the X-29A research program at the National Transonic Facility. This wind tunnel test leveraged the X-29A high alpha flight test program by enabling ground-to-flight correlation studies with an emphasis on Reynolds number effects. The background and objectives of this test program, as well as the comparison of high Reynolds number wind tunnel data to X-29A flight test data are presented. The effects of Reynolds number on the forebody pressures at high angles of attack are also presented. The purpose of this paper is to document this test and serve as a reference for future ground-to-flight correlation studies, and high angle-of-attack investigations. Good ground-to-flight correlations were observed for angles of attack up to 50 deg, and Reynolds number effects were also observed.
Lee, S. C.
The viscous effect on aerodynamic performance of an arbitrary airfoil executing low frequency maneuvers during transonic flight was investigated. The small disturbance code, LTRAN2, was modified by using a conventional integral method, BLAYER, for the boundary layer and an empirical relation, viscous wedge, for simulating the suddenly thickened boundary layer behind the shock. Before the shock, only the boundary layer displacement thickness was evaluated. After the shock, the empirical wedge thickness was superimposed on the boundary layer thickness along the surface as well as in the wake region. The pressure coefficients were calculated for both steady and unsteady states. The viscous solution takes fewer iterations to obtain the converged steady state solution. Comparisons made with experimental data and the inviscid solution show that the viscous solution agrees better with the experimental data with about the same (or slightly less) amount of computational time.
Kennelly, R. A., Jr.
An improved method for use of optimization techniques in transonic airfoil design is demonstrated. FLO6QNM incorporates a modified quasi-Newton optimization package, and is shown to be more reliable and efficient than the method developed previously at NASA-Ames, which used the COPES/CONMIN optimization program. The design codes are compared on a series of test cases with known solutions, and the effects of problem scaling, proximity of initial point to solution, and objective function precision are studied. In contrast to the older method, well-converged solutions are shown to be attainable in the context of engineering design using computational fluid dynamics tools, a new result. The improvements are due to better performance by the optimization routine and to the use of problem-adaptive finite difference step sizes for gradient evaluation.
Batill, Stephen M.
This report documents the results of a study which was conducted in order to establish a framework for the quantitative description of the uncertainty in measurements conducted in the National Transonic Facility (NTF). The importance of uncertainty analysis in both experiment planning and reporting results has grown significantly in the past few years. Various methodologies have been proposed and the engineering community appears to be 'converging' on certain accepted practices. The practical application of these methods to the complex wind tunnel testing environment at the NASA Langley Research Center was based upon terminology and methods established in the American National Standards Institute (ANSI) and the American Society of Mechanical Engineers (ASME) standards. The report overviews this methodology.
Gally, Thomas A.; Carlson, Leland A.
An inverse wing design method has been developed around an existing transonic wing analysis code. The original analysis code, TAWFIVE, has as its core the numerical potential flow solver, FLO30, developed by Jameson and Caughey. Features of the analysis code include a finite-volume formulation; wing and fuselage fitted, curvilinear grid mesh; and a viscous boundary layer correction that also accounts for viscous wake thickness and curvature. The development of the inverse methods as an extension of previous methods existing for design in Cartesian coordinates is presented. Results are shown for inviscid wing design cases in super-critical flow regimes. The test cases selected also demonstrate the versatility of the design method in designing an entire wing or discontinuous sections of a wing.
Walitt, L.; Liu, C. Y.
The viscous transonic flow about a stationary body in free air was numerically investigated. The geometry chosen was a symmetric NACA 64A010 airfoil at a freestream Mach number of 0.8, a Reynolds number of 4 million based on chord, and angles of attack of 0 and 2 degrees. These conditions were such that, at 2 degrees incidence unsteady periodic motion was calculated along the aft portion of the airfoil and in its wake. Although no unsteady measurements were made for the NACA 64A010 airfoil at these flow conditions, interpolated steady measurements of lift, drag, and surface static pressures compared favorably with corresponding computed time-averaged lift, drag, and surface static pressures.
Chang, J. F.; Lan, C. E.
A nonuniform transonic airfoil code is developed for applications in analysis, inverse design and direct optimization involving an airfoil immersed in propfan slipstream. Problems concerning the numerical stability, convergence, divergence and solution oscillations are discussed. The code is validated by comparing with some known results in incompressible flow. A parametric investigation indicates that the airfoil lift-drag ratio can be increased by decreasing the thickness ratio. A better performance can be achieved if the airfoil is located below the slipstream center. Airfoil characteristics designed by the inverse method and a direct optimization are compared. The airfoil designed with the method of direct optimization exhibits better characteristics and achieves a gain of 22 percent in lift-drag ratio with a reduction of 4 percent in thickness.
Liou, M.-S.; Coakley, T. J.
Forced and naturally occurring, self-sustaining oscillations of transonic flows in two-dimensional diffusers were computed using MacCormack's hybrid method. Depending upon the shock strengths and the area ratios, the flow was fully attached or separated by either the shock or the adverse pressure gradient associated with the enlarging diffuser area. In the case of forced oscillations, a sinusoidal plane pressure wave at frequency 300 Hz was prescribed at the exit. A sufficiently large amount of data were acquired and Fourier analyzed. The distrbutions of time-mean pressures, the power spectral density, and the amplitude with phase angle along the top wall and in the core region were determined. Comparison with experimental results for the forced oscillation generally gave very good agreement; some success was achieved for the case of self-sustaining oscillation despite substantial three-dimensionality in the test. An observation of the sequence of self-sustaining oscillations was given.
Handa, T.; Miyazato, Y.; Masuda, M.; Matsuo, K.
Multiple shocklets are frequently generated in transonic diffuser flows. The present paper investigates the formation of these shocklets with a high-speed CCD camera combined with the schlieren method. It is observed that compression waves steepen while propagating upstream, and eventually become new shock waves. The ordinary shock wave is found to move upstream beyond the nozzle throat or to disappear while moving downstream depending on the pressure ratio across the nozzle. This phenomenon is also analyzed with the one-dimensional Euler equations by assuming a pressure disturbance given by the sine function at the channel exit. The calculated results are found to reproduce quite well the experimental behavior of the shocklets. The effect of the frequency of disturbance is also studied numerically, and it is shown that the multiple shocklet pattern appears when the amplitude of disturbance is not large and the diverging part of the channel downstream of the ordinary shock wave is long.
Burner, A. W.; Snow, W. L.; Goad, W. K.
A photogrammetric closed circuit television system to measure model deformation at the National Transonic Facility is described. The photogrammetric approach was chosen because of its inherent rapid data recording of the entire object field. Video cameras are used to acquire data instead of film cameras due to the inaccessibility of cameras which must be housed within the cryogenic, high pressure plenum of this facility. A rudimentary theory section is followed by a description of the video-based system and control measures required to protect cameras from the hostile environment. Preliminary results obtained with the same camera placement as planned for NTF are presented and plans for facility testing with a specially designed test wing are discussed.
Crawley, E. F.
A method was developed and demonstrated for the direct measurement of aerodynamic forcing and aerodynamic damping of a transonic compressor. The method is based on the inverse solution of the structural dynamic equations of motion of the blade disk system in order to determine the forces acting on the system. The disturbing and damping forces acting on a given blade are determined if the equations of motion are expressed in individual blade coordinates. If the structural dynamic equations are transformed to multiblade coordinates, the damping can be measured for blade disk modes, and related to a reduced frequency and interblade phase angle. In order to measure the aerodynamic damping in this way, the free response to a known excitation is studied.
Malone, Michael B.; Peavey, Charles C.
The Transonic Nozzle Boattail Drag Study was initiated in 1995 to develop an understanding of how external nozzle transonic aerodynamics effect airplane performance and how strongly those effects are dependent on nozzle configuration (2D vs. axisymmetric). MDC analyzed the axisymmetric nozzle. Boeing subcontracted Northrop-Grumman to analyze the 2D nozzle. AU participants analyzed the AGARD nozzle as a check-out and validation case. Once the codes were checked out and the gridding resolution necessary for modeling the separated flow in this region determined, the analysis moved to the installed wing/body/nacelle/diverter cases. The boat tail drag validation case was the AGARD B.4 rectangular nozzle. This test case offered both test data and previous CFD analyses for comparison. Results were obtained for test cases B.4.1 (M=0.6) and B.4.2 (M=0.938) and compared very well with the experimental data. Once the validation was complete a CFD grid was constructed for the full Ref. H configuration (wing/body/nacelle/diverter) using a combination of patched and overlapped (Chimera) grids. This was done to ensure that the grid topologies and density would be adequate for the full model. The use of overlapped grids allowed the same grids from the full configuration model to be used for the wing/body alone cases, thus eliminating the risk of grid differences affecting the determination of the installation effects. Once the full configuration model was run and deemed to be suitable the nacelle/diverter grids were removed and the wing/body analysis performed. Reference H wing/body results were completed for M=0.9 (a=0.0, 2.0, 4.0, 6.0 and 8.0), M=1.1 (a=4.0 and 6.0) and M=2.4 (a=0.0, 2.0, 4.4, 6.0 and 8.0). Comparisons of the M=0.9 and M=2.4 cases were made with available wind tunnel data and overall comparisons were good. The axi-inlet/2D nozzle nacelle was analyzed isolated. The isolated nacelle data coupled with the wing/body result enabled the interference effects of the
Boppe, C. W.
A computational method for simulating the aerodynamics of wing-fuselage configurations at transonic speeds is developed. The finite difference scheme is characterized by a multiple embedded mesh system coupled with a modified or extended small disturbance flow equation. This approach permits a high degree of computational resolution in addition to coordinate system flexibility for treating complex realistic aircraft shapes. To augment the analysis method and permit applications to a wide range of practical engineering design problems, an arbitrary fuselage geometry modeling system is incorporated as well as methodology for computing wing viscous effects. Configuration drag is broken down into its friction, wave, and lift induced components. Typical computed results for isolated bodies, isolated wings, and wing-body combinations are presented. The results are correlated with experimental data. A computer code which employs this methodology is described.
Heeg, Jennifer; Chwalowski, Pawel
This paper builds on the computational aeroelastic results published previously and generated in support of the second Aeroelastic Prediction Workshop for the NASA Benchmark Supercritical Wing configuration. The computational results are obtained using FUN3D, an unstructured grid Reynolds-Averaged Navier-Stokes solver developed at the NASA Langley Research Center. The analysis results focus on understanding the dip in the transonic flutter boundary at a single Mach number (0.74), exploring an angle of attack range of ??1 to 8 and dynamic pressures from wind off to beyond flutter onset. The rigid analysis results are examined for insights into the behavior of the aeroelastic system. Both static and dynamic aeroelastic simulation results are also examined.
Corliss, James M.; Cole, Stanley, R.
The heavy gas test medium has recently been changed in the Transonic Dynamics Tunnel (TDT) at the NASA Langley Research Center. A NASA Construction of Facilities project has converted the TDT heavy gas from dichlorodifluoromethane (R12) to 1,1,1,2 tetrafluoroethane (R134a). The facility s heavy gas processing system was extensively modified to implement the conversion to R134a. Additional system modifications have improved operator interfaces, hardware reliability, and quality of the research data. The facility modifications included improvements to the heavy gas compressor and piping, the cryogenic heavy gas reclamation system, and the heavy gas control room. A series of wind tunnel characterization and calibration tests are underway. Results of the flow characterization tests show the TDT operating envelope in R134a to be very similar to the previous operating envelope in R12.
Lee, In; Kim, Jong-Yun; Kim, Kyung-Seok; Lim, In-Gyu
Flight vehicles experience aeroelastic problems due to the interaction between structures and aerodynamic forces. Aeroelastic instability is usually a critical problem in transonic and lower supersonic regions. In present study, the aeroelastic analyses of several flight vehicles have been performed using the coupled techniques of computational fluid dynamics (CFD) and computational structural dynamics (CSD). The aeroelastic characteristics based on several aircraft models are investigated using the developed aeroelastic analysis system. On the other hand, structural nonlinearities always exist in flight vehicles. Structural nonlinearities such as freeplay and large deformation effects are considered in the present aeroelastic analysis system. Finally, aeroelastic characteristics of several flight vehicles will be explained considering both aerodynamic and structural nonlinearities.
Gloss, Blair B.; Bruce, Robert A.
As cryogenic wind tunnels are utilized, problems associated with the low temperature environment are being discovered and solved. Recently, water vapor contamination was discovered in the National Transonic Facility, and the source was shown to be the internal insulation which is a closed-cell polyisocyanurate foam. After an extensive study of the absorptivity characteristics of the NTF thermal insulation, the most practical solution to the problem was shown to be the maintaining of a dry environment in the circuit at all times. Utilizing a high aspect ratio transport model, it was shown that the moisture contamination effects on the supercritical wing pressure distributions were within the accuracy of setting test conditions and as such were considered negligible for this model.
Verdon, J. M.; Caspar, J. R.
A potential flow analysis to predict unsteady airloads produced by the vibrations of turbomachinery blades operating at transonic Mach numbers is presented. The unsteady aerodynamic model includes the effects of blade geometry, finite mean pressure variation across the blade row, high frequency blade motion, and shock motion within the framework of a linearized, frequency domain formulation. The unsteady equations are solved implicit, least squares, finite difference approximation which is applicable on arbitrary grids. A numerical solution for the entire unsteady field is determined by matching a solution determined on a rectilinear type cascade mesh, which covers an extended blade passage region, to a solution determined on a detailed polar type local mesh, which covers and extends well beyond the supersonic region(s) adjacent to a blade surface. Cascades of double circular arc and flat plate blades demonstrate the unsteady analysis, and partially illustrate the effects of blade geometry, inlet Mach number, blade vibration frequency and shock motion on unsteady response.
Burner, Alpheus W.; Wahls, Richard A.; Goad, William K.
A technique for measuring wing twist currently in use at the National Transonic Facility is described. The technique is based upon a single camera photogrammetric determination of two dimensional coordinates with a fixed (and known) third dimensional coordinate. The wing twist is found from a conformal transformation between wind-on and wind-off 2-D coordinates in the plane of rotation. The advantages and limitations of the technique as well as the rationale for selection of this particular technique are discussed. Examples are presented to illustrate run-to-run and test-to-test repeatability of the technique in air mode. Examples of wing twist in cryogenic nitrogen mode are also presented.
Gertz, J. B.; Epstein, A. H.
High-frequency response probes which had previously been used exclusively in the MIT Blowndown Facility were successfully employed in two conventional steady state axial flow compressor facilities to investigate the unsteady flowfields of highly loaded transonic compressors at design point operation. Laser anemometry measurements taken simultaneously with the high response data were also analyzed. The time averaged high response data of static and total pressure agreed quite well with the conventional steady state instrumentation except for flow angle which showed a large spread in values at all radii regardless of the type of instrumentation used. In addition, the time resolved measurements confirmed earlier test results obtained in the MIT Blowdown Facility for the same compressor. The results of these tests have further revealed that the flowfields of highly loaded transonic compressors are heavily influenced by unsteady flow phenomena. The high response measurements exhibited large variations in the blade to blade flow and in the blade passage flow. The observed unsteadiness in the blade wakes is explained in terms of the rotor blades' shed vorticity in periodic vortex streets. The wakes were modeled as two-dimensional vortex streets with finite size cores. The model fit the data quite well as it was able to reproduce the average wake shape and bi-modal probability density distributions seen in the laser anemometry data. The presence of vortex streets in the blade wakes also explains the large blade to blade fluctuations seen by the high response probes which is simply due to the intermittent sampling of the vortex street as it is swept past a stationary probe.
Chima, Rodrick V.
One goal of the NASA Fundamental Aeronautics Program is the assessment of computational fluid dynamic (CFD) codes used for the design and analysis of many aerospace systems. This paper describes the assessment of the SWIFT turbomachinery analysis code for two similar transonic compressors, NASA rotor 37 and stage 35. The two rotors have identical blade profiles on the front, transonic half of the blade but rotor 37 has more camber aft of the shock. Thus the two rotors have the same shock structure and choking flow but rotor 37 produces a higher pressure ratio. The two compressors and experimental data are described here briefly. Rotor 37 was also used for test cases organized by ASME, IGTI, and AGARD in 1994-1998. Most of the participating codes over predicted pressure and temperature ratios, and failed to predict certain features of the downstream flowfield. Since then the AUSM+ upwind scheme and the k- turbulence model have been added to SWIFT. In this work the new capabilities were assessed for the two compressors. Comparisons were made with overall performance maps and spanwise profiles of several aerodynamic parameters. The results for rotor 37 were in much better agreement with the experimental data than the original blind test case results although there were still some discrepancies. The results for stage 35 were in very good agreement with the data. The results for rotor 37 were very sensitive to turbulence model parameters but the results for stage 35 were not. Comparison of the rotor solutions showed that the main difference between the two rotors was not blade camber as expected, but shock/boundary layer interaction on the casing.
Lepicovsky, J.; Capece, V. R.; Ford, C. T.
Investigations of unsteady pressure loadings on the blades of fans operating near the stall flutter boundary are carried out under simulated conditions in the NASA Transonic Flutter Cascade facility (TFC). It has been observed that for inlet Mach numbers of about 0.8, the cascade flowfield exhibits intense low-frequency pressure oscillations. The origins of these oscillations were not clear. It was speculated that this behavior was either caused by instabilities in the blade separated flow zone or that it was a tunnel resonance phenomenon. It has now been determined that the strong low-frequency oscillations, observed in the TFC facility, are not a cascade phenomenon contributing to blade flutter, but that they are solely caused by the tunnel resonance characteristics. Most likely, the self-induced oscillations originate in the system of exit duct resonators. For sure, the self-induced oscillations can be significantly suppressed for a narrow range of inlet Mach numbers by tuning one of the resonators. A considerable amount of flutter simulation data has been acquired in this facility to date, and therefore it is of interest to know how much this tunnel self-induced flow oscillation influences the experimental data at high subsonic Mach numbers since this facility is being used to simulate flutter in transonic fans. In short, can this body of experimental data still be used reliably to verify computer codes for blade flutter and blade life predictions? To answer this question a study on resonance effects in the NASA TFC facility was carried out. The results, based on spectral and ensemble averaging analysis of the cascade data, showed that the interaction between self-induced oscillations and forced blade motion oscillations is very weak and can generally be neglected. The forced motion data acquired with the mistuned tunnel, when strong self-induced oscillations were present, can be used as reliable forced pressure fluctuations provided that they are extracted
Zaman, Khairul B. M. Q.
Noise characteristics of convergent-divergent (C-D) nozzles in the overexpanded regime are the focus of this paper. The flow regime is encountered during takeoff and landing of certain airplanes and also with rocket nozzles in launch-pad environment. Experimental results from laboratory-scale single nozzles are discussed. The flow often undergoes a resonance accompanied by emission of tones (referred to as transonic tones). The phenomenon is different from the well-known screech tones. Unlike screech, the frequency increases with increasing supply pressure. There is a staging behavior odd harmonic stages occur at lower pressures while the fundamental occurs in a range of relatively higher pressures. A striking feature is that tripping of the nozzle s internal boundary layer tends to suppress the resonance. However, even in the absence of tones the broadband levels are found to be high. That is, relative to a convergent case and at same pressure ratio, the C-D nozzles are found to be noisier, often by more than 10dB. This excess broadband noise (referred to as EBBN) is further explored. Its characteristics are found to be different from the well-known broadband shockassociated noise ( BBSN ). For example, while the frequency of the BBSN peak varies with observation angle no such variation is noted with EBBN. The mechanisms of the transonic tone and the EBBN are not completely understood yet. They appear to be due to unsteady shock motion inside the nozzle. The shock drives the flow downstream like a vibrating diaphragm, and resonance takes place similarly as with acoustic resonance of a conical section having one end closed and the other end open. When the boundary layer is tripped, apparently a breakdown of azimuthal coherence suppresses the resonance. However, there is still unsteady shock motion albeit with superimposed randomness. Such random motion of the internal shock and its interaction with the separated boundary layer produces the EBBN.
Bryanston-Cross, Peter J.; Burnett, Mark; Lee, Wing K. A.; Udrea, Doina D.; Chana, Kamaljit S.; Anderson, S. J.
A series of measurement have been made using PIV in the trailing edge region of the stator row and rotor in the annular transonic cascade at RAe Farnborough. The measurements provide an instantaneous quantitative whole field visualization of an unsteady transonic flow interaction region. This work is the first such measurement to be made in a rotating transonic facility.
Ecer, A.; Akay, H. U.
The finite element method is applied for the solution of transonic potential flows through a cascade of airfoils. Convergence characteristics of the solution scheme are discussed. Accuracy of the numerical solutions is investigated for various flow regions in the transonic flow configuration. The design of an efficient finite element computational grid is discussed for improving accuracy and convergence.
Garriz, Javier A.; Haigler, Kara J.
A three dimensional transonic Wind-tunnel Interference Assessment and Correction (WIAC) procedure developed specifically for use in the National Transonic Facility (NTF) at NASA Langley Research Center is discussed. This report is a user manual for the codes comprising the correction procedure. It also includes listings of sample procedures and input files for running a sample case and plotting the results.
Edwards, John W.
A viscous-inviscid interactive coupling method is used for the computation of unsteady transonic flows involving separation and reattachment. A lag-entrainment integral boundary layer method is used with the transonic small disturbance potential equation in the CAP-TSDV (Computational Aeroelasticity Program - Transonic Small Disturbance) code. Efficient and robust computations of steady and unsteady separated flows, including steady separation bubbles and self-excited shock-induced oscillations are presented. The buffet onset boundary for the NACA 0012 airfoil is accurately predicted and shown computationally to be a Hopf bifurcation. Shock-induced oscillations are also presented for the 18 percent circular arc airfoil. The oscillation onset boundaries and frequencies are accurately predicted, as is the experimentally observed hysteresis of the oscillations with Mach number. This latter stability boundary is identified as a jump phenomenon. Transonic wing flutter boundaries are also shown for a thin swept wing and for a typical business jet wing, illustrating viscous effects on flutter and the effect of separation onset on the wing response at flutter. Calculations for both wings show limit cycle oscillations at transonic speeds in the vicinity of minimum flutter speed indices.
Rivers, S. M. B.; Wahls, R. A.
Results of turbulence model comparisons from two studies on supersonic transport configurations performed during the NASA High-speed Research program are given. Results are presented for both transonic conditions at Mach 0.90 and supersonic conditions at Mach 2.48. A feature of these two studies was the availability of higher Reynolds number wind tunnel data with which to compare the computational results. The transonic wind tunnel data was obtained in the National Transonic Facility at NASA Langley, and the supersonic data was obtained in the Boeing Polysonic Wind Tunnel. The computational data was acquired using a state of the art Navier-Stokes flow solver with a wide range of turbulence models implemented. The results show that the computed forces compare reasonably well with the experimental data, with the Baldwin- Lomax with Degani-Schiff modifications and the Baldwin-Barth models showing the best agreement for the transonic conditions and the Spalart-Allmaras model showing the best agreement for the supersonic conditions. The transonic results were more sensitive to the choice of turbulence model than were the supersonic results.
Bledsoe, Margaret Randolph
We calculate shockless transonic flows past two -dimensional cascades of airfoils characterized by a prescribed speed distribution. The approach is to find solutions of the partial differential equation (c('2)-u('2)) (PHI)(,xx) - 2uv (PHI)(,xy) + (c('2)-v('2)) (PHI)(,yy) = 0 by the method of complex characteristics. Here (PHI) is the velocity potential, so (DEL)(PHI) = (u,v), and c is the local speed of sound. Our method consists in noting that the coefficients of the equation are analytic, so that we can use analytic continuation, conformal mapping, and a spectral method in the hodograph plane to determine the flow. After complex extension we obtain canonical equations for (PHI) and for the stream function (psi) as well as an explicit map from the hodograph plane to complex characteristic coordinates. In the subsonic case, a new coordinate system is defined in which the flow region corresponds to the interior of an ellipse. We construct special solutions of the flow equations in these coordinates by solving characteristic initial value problems in the ellipse with initial data defined by the complete system of Chebyshev polynomials. The condition (psi) = 0 on the boundary of the ellipse is used to determine the series representation of (PHI) and (psi). The map from the ellipse to the complex flow coordinates is found from data specifying the speed q as a function of the arc length s. The transonic problem for shockless flow becomes well posed after appropriate modifications of this procedure. The nonlinearity of the problem is handled by an iterative method that determines the boundary value problem in the ellipse and the map function in sequence. We have implemented this method as a computer code to design two-dimensional cascades of shockless compressor airfoils with gap-to-chord ratios as low as .5 and supersonic zones on both the upper and lower surfaces. The method may be extended to solve more general boundary value problems for second order partial
Atwood, C. A.
Numerical simulations of two classes of unsteady flows are obtained via the Navier-Stokes equations: a blast-wave/target interaction problem class and a transonic cavity flow problem class. The method developed for the viscous blast-wave/target interaction problem assumes a laminar, perfect gas implemented in a structured finite-volume framework. The approximately factored implicit scheme uses Newton subiterations to obtain the spatially and temporally second-order accurate time history of the blast-waves with stationary targets. The inviscid flux is evaluated using either of two upwind techniques, while the full viscous terms are computed by central differencing. Comparisons of unsteady numerical, analytical, and experimental results are made in two- and three-dimensions for Couette flows, a starting shock-tunnel, and a shock-tube blockage study. The results show accurate wave speed resolution and nonoscillatory discontinuity capturing of the predominantly inviscid flows. Viscous effects were increasingly significant at large post-interaction times. While the blast-wave/target interaction problem benefits from high-resolution methods applied to the Euler terms, the transonic cavity flow problem requires the use of an efficient scheme implemented in a geometrically flexible overset mesh environment. Hence, the Reynolds averaged Navier-Stokes equations implemented in a diagonal form are applied to the cavity flow class of problems. Comparisons between numerical and experimental results are made in two-dimensions for free shear layers and both rectangular and quieted cavities, and in three-dimensions for Stratospheric Observatory For Infrared Astronomy (SOFIA) geometries. The acoustic behavior of the rectangular and three-dimensional cavity flows compare well with experiment in terms of frequency, magnitude, and quieting trends. However, there is a more rapid decrease in computed acoustic energy with frequency than observed experimentally owing to numerical
Giel, P. W.; Thurman, D. R.; VanFossen, G. J.; Hippensteele, S. A.; Boyle, R. J.
Turbine blade endwall heat transfer measurements are given for a range of Reynolds and Mach numbers. Data were obtained for Reynolds numbers based on inlet conditions of 0.5 and 1.0 x 106, for isentropic exit Mach numbers of 1.0 and 1.3, and for freestream turbulence intensities of 0.25% and 7.0%. Tests were conducted in a linear cascade at the NASA Lewis Transonic Turbine Blade Cascade Facility. The test article was a turbine rotor with 136' of turning and an axial chord of 12.7 cm. The large scale allowed for very detailed measurements of both flow field and surface phenomena. The intent of the work is to provide benchmark quality data for computational fluid dynamics (CFD) code and model verification. The flow field in the cascade is highly three-dimensional as a result of thick boundary layers at the test section inlet. Endwall heat transfer data were obtained using a steady-state liquid crystal technique.
Munson, Matthew; Rempfer, Dietmar; Williams, David; Acharya, Mukund
The ability to control flow separation angles from compressor inlet guide vanes with a Coanda-type actuator is demonstrated using both wind tunnel experiments and finite element simulations. Vectoring angles up to 40 degrees from the uncontrolled baseline state were measured with helium schlieren visualization at transonic Mach numbers ranging from 0.1 to 0.6, and with airfoil chord Reynolds numbers ranging from 89,000 to 710,000. The magnitude of the vectoring angle is shown to depend upon the geometry of the trailing edge, and actuator slot size, and the momentum flux coefficient. Under certain conditions the blowing has no effect on the vectoring angle indicating that the Coanda effect is not present. DNS simulations with the finite element method investigated the effects of geometry changes and external flow. Continuous control of the vectoring angle is demonstrated, which has important implications for application to rotating machinery. The technique is shown to reduce the stall flow coefficient by 15 percent in an axial flow compressor.
Chima, R. V.
The flow through the tip clearance region of a transonic compressor rotor (NASA rotor 37) was computed and compared to aerodynamic probe and laser anemometer data. Tip clearance effects were modeled both by gridding the clearance gap and by using a simple periodicity model across the ungridded gap. The simple model was run with both the full gap height, and with half the gap height to simulate a vena-contracta effect. Comparisons between computed and measured performance maps and downstream profiles were used to validate the models and to assess the effects of gap height on the simple clearance model. Recommendations were made concern- ing the use of the simple clearance model Detailed comparisons were made between the gridded clearance gap solution and the laser anemometer data near the tip at two operating points. The computed results agreed fairly well with the data but overpredicted the extent of the casing separation and underpredicted the wake decay rate. The computations were then used to describe the interaction of the tip vortex, the passage shock, and the casing boundary layer.
Chima, R. V.
The flow through the tip clearance region of a transonic compressor rotor (NASA rotor 37) was computed and compared to aerodynamic probe and laser anemometer data. Tip clearance effects were modeled both by gridding the clearance gap and by using a simple periodicity model across the ungridded gap. The simple model was run with both the full gap height, and with half the gap height to simulate a vena-contracta effect. Comparisons between computed and measured performance maps and downstream profiles were used to validate the models and to assess the effects of gap height on the simple clearance model. Recommendations were made concerning the use of the simple clearance model. Detailed comparisons were made between the gridded clearance gap solution and the laser anemometer data near the tip at two operating points. The computer results agreed fairly well with the data but overpredicted the extent of the casing separation and underpredicted the wake decay rate. The computations were then used to describe the interaction of the tip vortex, the passage shock, and the casing boundary layer.
Yetter, J. A.; Abeyounis, W. K.
A flow survey has been made of the test section of the NASA Langley Research Center 16-Foot Transonic Tunnel at subsonic and supersonic speeds. The survey was performed using five five-hole pyramid-head probes mounted at 14 inch intervals on a survey rake. Probes were calibrated at freestream Mach numbers from 0.50 to 0.95 and from 1.18 to 1.23. Flowfield surveys were made at Mach numbers from 0.50 to 0.90 and at Mach 1.20. The surveys were made at tunnel stations 130.6, 133.6, and 136.0. By rotating the survey rake through 180 degrees, a cylindrical volume of the test section 4.7 feet in diameter and 5.4 feet long centered about the tunnel centerline was surveyed. Survey results showing the measured test section upflow and sideflow characteristics and local Mach number distributions are presented. The report documents the survey probe calibration techniques used, summarizes the procedural problems encountered during testing, and identifies the data discrepancies observed during the post-test data analysis.
Tauber, M. E.
The high noise levels produced by helicopters are major sources of concern. There are many sources of the noise, but during high-speed forward flight, impulsive noise dominates the noise spectrum. The cause of the high-speed impulsive noise is the propagation into the far field of shock waves that form on the advancing blade. This mechanism has been labeled 'delocalization'. It has been shown, however, that by judicious design of the blade-tip planform, delocalization can be prevented. The objective of the present study is to illustrate how blade-tip configurations (both planform and airfoil shape) can be systematically varied to identify shapes that avoid delocalization and simultaneously improve aerodynamic performance. This has been done using the latest version of the ROT22 transonic, full-potential, quasi-steady, rotor flow-field code. A hypothetical modern rotor blade was postulated, and tip modifications consisting of taper, sweep, and airfoil section alterations were investigated. Planform modifications were found to be most effective in eliminating delocalization.
Wong, Y. S.
A computational technique for the solution of the full potential equation is presented. The method consists of outer and inner iterations. The outer iterate is based on a Newton like algorithm, and a preconditioned Minimal Residual method is used to seek an approximate solution of the system of linear equations arising at each inner iterate. The present iterative scheme is formulated so that the uncertainties and difficulties associated with many iterative techniques, namely the requirements of acceleration parameters and the treatment of additional boundary conditions for the intermediate variables, are eliminated. Numerical experiments based on the new method for transonic potential flows around the NACA 0012 airfoil at different Mach numbers and different angles of attack are presented, and these results are compared with those obtained by the Approximate Factorization technique. Extention to three dimensional flow calculations and application in finite element methods for fluid dynamics problems by the present method are also discussed. The Inexact Newton like method produces a smoother reduction in the residual norm, and the number of supersonic points and circulations are rapidly established as the number of iterations is increased.
Aoyama, Takashi; Saito, Shigeru
Three-dimensional Euler equations are directly solved to analyze the High-Speed Impulsive (HSI) noise of a helicopter motor by using CFD technique. The MSI noise is one of the most important sources of helicopter noise. It is generated on the advancing side of a helicopter caused by the shock wave on a blade surface. Although the method which solves the Ffowcs-Williams and Hawkings equation has been often used to analyze the subsonic rotor noise, it doesn't success to predict the transonic rotor noise such as the HSI noise With the advance of CFD technique, the calculation of the HSI noise is recently performed by the combined method of CFD with the Kirchhoff's equation or by the direct simulation using CFD technique. The latter has not been studied enough because huge number of grids are needed to capture the propagation of sound from a blade to an observer located in a far field. So, the powerful supercomputer of NAL, Numerical Wind Tunnel (NWT) is employed to calculate the RSI noise of a non-lifting hovering rotor directly by using the method. The numerical method to solve the governing equation is an implicit finite-difference scheme which utilizes a higher-order upwind scheme based on TVD. As a result, it is observed that the calculated wave form is in very good agreement with an experimental data at sonic cylinder. The agreement is not very good at about three rotor radii but is reasonable at about two rotor radii.
Latham, E.; Gawienowski, J.; Meriwether, F.
A 15 percent scale lightweight fighter type inlet forebody was tested in the Ames 14 foot transonic wind tunnel at Mach numbers of 0.7, 0.9, and 1.04. The inlet was a two dimensional horizontal ramp system designed for a Mach number of 2.2. Four inlet devices designed to prevent or delay cowl-lip boundary layer separation or to improve the inlet internal flow characteristics at high angles of attack were investigated. The devices used to control cowl-lip separation consisted of cowl leading edge flaps, slotted flaps, and tangential blowing. To improve the internal flow characteristics, discrete jet nozzle flows were directed downstream and parallel to the duct surface in the subsonic diffuser to energize the wall boundary layer. The discrete jets used in the subsonic diffuser were also tested in combination with each of the cowl leading edge devices. Test measurements included engine-face total pressure recovery, steady state distortion, dynamic distortion, duct boundary layer profiles, and duct-surface static pressures.
The present work was motivated by an ongoing research program at NASA Langley Research Center to develop a semi-span testing capability for the National Transonic Facility (NTF). This test technique is being investigated as a means to design and optimize high-lift devices at flight Reynolds numbers in a ground test facility. Even though the freestream Mach numbers of interest are around .20, the flow around a transport wing with high lift devices deployed may contain regions of compressible flow. Thus to properly model the flow physics, a compressible flow solver may be required. However, the application of a compressible flow solver at low Mach numbers can be problematic. The objective of this phase of the project is to directly compare the performance of two widely used three-dimensional compressible Navier-Stokes solvers at low Mach numbers to both experimental data and to results obtained from an incompressible Navier-Stokes solver. The geometries of interest are two isolated wings with different leading edge sweep angles. The compressible Navier-Stokes solvers chosen, TLNS3D-MB and CFL3D, which were developed at NASA Langley Research Center (LaRC), represent the current state-of-the-art in compressible 3-D Navier-Stokes solvers. The incompressible Navier-Stokes solver, INS3D-UP, developed recently at NASA Ames Research Center (ARC), represents the current state-of-the-art in incompressible Navier-Stokes solvers.
Walter, Joel A.; Lawrence, William R.; Elder, David W.; Treece, Michael D.
This paper introduces an uncertainty model being developed for the National Transonic Facility (NTF). The model uses a Monte Carlo technique to propagate standard uncertainties of measured values through the NTF data reduction equations to calculate the combined uncertainties of the key aerodynamic force and moment coefficients and freestream properties. The uncertainty propagation approach to assessing data variability is compared with ongoing data quality assessment activities at the NTF, notably check standard testing using statistical process control (SPC) techniques. It is shown that the two approaches are complementary and both are necessary tools for data quality assessment and improvement activities. The SPC approach is the final arbiter of variability in a facility. Its result encompasses variation due to people, processes, test equipment, and test article. The uncertainty propagation approach is limited mainly to the data reduction process. However, it is useful because it helps to assess the causes of variability seen in the data and consequently provides a basis for improvement. For example, it is shown that Mach number random uncertainty is dominated by static pressure variation over most of the dynamic pressure range tested. However, the random uncertainty in the drag coefficient is generally dominated by axial and normal force uncertainty with much less contribution from freestream conditions.
Piatak, David J.; Cleckner, Craig S.
A new forced oscillation system has been installed and tested at NASA Langley Research Center's Transonic Dynamics Tunnel (TDT). The system is known as the Oscillating Turntable (OTT) and has been designed for the purpose of oscillating, large semispan models in pitch at frequencies up to 40 Hz to acquire high-quality unsteady pressure and loads data. Precisely controlled motions of a wind-tunnel model on the OTT can yield unsteady aerodynamic phenomena associated with flutter, limit cycle oscillations, shock dynamics, and non-linear aerodynamic effects on many vehicle configurations. This paper will discuss general design and components of the OTT and will present test data from performance testing and from research tests on two rigid semispan wind-tunnel models. The research tests were designed to challenge the OTT over a wide range of operating conditions while acquiring unsteady pressure data on a small rectangular supercritical wing and a large supersonic transport wing. These results will be presented to illustrate the performance capabilities, consistency of oscillations, and usefulness of the OTT as a research tool.
Liu, Yangwei; Wu, Jianuo; Lu, Lipeng
Eight turbulence models frequently used in aerodynamics have been employed in the detailed numerical investigations for transonic flows in the Sajben diffuser, to assess the predictive capabilities of the turbulence models for shock wave/turbulent boundary layer interactions (SWTBLI) in internal flows. The eight turbulence models include: the Spalart-Allmaras model, the standard k - 𝜀 model, the RNG k - 𝜀 model, the realizable k - 𝜀 model, the standard k - ω model, the SST k - ω model, the v2¯ - f model and the Reynolds stress model. The performance of the different turbulence models adopted has been systematically assessed by comparing the numerical results with the available experimental data. The comparisons show that the predictive performance becomes worse as the shock wave becomes stronger. The v2¯ - f model and the SST k - ω model perform much better than other models, and the SST k - ω model predicts a little better than the v2¯ - f model for pressure on walls and velocity profile, whereas the v2¯ - f model predicts a little better than the SST k - ω model for separation location, reattachment location and separation length for strong shock case.
Gissen, Abraham N.; Vukasinovic, Bojan; Glezer, Ari; Gogineni, Sivaram P.
The effects of fluidic actuation on the evolution and dynamics of a transonic shock over a two-dimensional convex surface by controlling the ensuing shock-induced separation are investigated in wind tunnel experiments. Actuation is effected by a spanwise array of high-frequency (nominally 10 kHz) fluidic oscillating jets. The flow field upstream and downstream of the shock is investigated using high-speed Schlieren and PIV (3,000fps), and surface pressure measurements. It is shown that control of the shock-induced separating shear layer by exploiting direct control of small-scale motion can alter the degree of flow attachment and have a profound effect on the shock dynamics. The actuation diminishes shock oscillations near the surface, and leads to streamwise shock displacement that is proportional to the actuation strength (as measured, for example, by the mass flow rate coefficient). The strong correlation between the shock displacement and surface pressure are explored for application of closed-loop control.
Reid, Lonnie; Celestina, Mark L.; Dewitt, Kenneth; Keith, Theo
This study involves an experimental and numerical investigation of the three-dimensional flows in a transonic compressor rotor. A variety of data which could be used, in a complementary fashion, to validate/calibrate the computational fluid dynamics turbomachinery code and improve understanding of the flow physics, were acquired. Detailed radial survey data which consisted of total pressure, total temperature, static pressure and flow angle were obtained at stations upstream and downstream of the rotor blade. Detailed velocity and turbulence profiles were obtained upstream of the rotor and used as the upstream boundary conditions for the numerical analysis. Calibrated flush-mounted hot film probes were used to measure wall shear stress on the hub and casing walls upstream of the rotor. The blade-to-blade shear-stress angle distributions were obtained at two axial locations on the rotor casing, using flush-mounted hot film probes. A numerical analysis conducted using a three-dimensional Navier-Stokes code was compared with the experimental results.
Brooks, Cuyler W., Jr.; Harris, Charles D.; Reagon, Patricia G.
The NASA Langley 8-Foot Transonic Pressure Tunnel is a continuous-flow, variable-pressure wind tunnel with control capability to independently vary Mach number, stagnation pressure, stagnation temperature, and humidity. The top and bottom walls of the test section are axially slotted to permit continuous variation of the test section Mach number from 0.2 to 1.2, the slot-width contour provides a gradient-free test section 50 in. long for Mach numbers equal to or greater than 1.0 and 100 in. long for Mach numbers less than 1.0. The stagnation pressure may be varied from 0.25 to 2.0 atm. The tunnel test section has been recalibrated to determine the relationship between the free-stream Mach number and the test chamber reference Mach number. The hardware was the same as that of an earlier calibration in 1972 but the pressure measurement instrumentation available for the recalibration was about an order of magnitude more precise. The principal result of the recalibration was a slightly different schedule of reentry flap settings for Mach numbers from 0.80 to 1.05 than that determined during the 1972 calibration. Detailed tunnel contraction geometry, test section geometry, and limited test section wall boundary layer data are presented.
Mineck, Raymond E.; Owens, Lewis R., Jr.; Wahls, Richard A.; Hannon, Judith A.
A computer model has been developed to simulate the processes involved in the operation of the National Transonic Facility (NTF), a large cryogenic wind tunnel at the Langley Research Center. The simulation was verified by comparing the simulated results with previously acquired data from three experimental wind tunnel test programs in the NTF. The comparisons suggest that the computer model simulates reasonably well the processes that determine the liquid nitrogen (LN2) consumption, electrical consumption, fan-on time, and the test time required to complete a test plan at the NTF. From these limited comparisons, it appears that the results from the simulation model are generally within about 10 percent of the actual NTF test results. The use of actual data acquisition times in the simulation produced better estimates of the LN2 usage, as expected. Additional comparisons are needed to refine the model constants. The model will typically produce optimistic results since the times and rates included in the model are typically the optimum values. Any deviation from the optimum values will lead to longer times or increased LN2 and electrical consumption for the proposed test plan. Computer code operating instructions and listings of sample input and output files have been included.
Wells, Douglas P.
NASA's investment and research in aviation has led to new technologies and concepts that make aircraft more efficient and environmentally friendly. One aircraft design operational concept is the reduction of cruise speed to reduce fuel burned during a mission. Although this is not a new idea, it was used by all of the contractors involved in a 2008 NASA sponsored study that solicited concept and technology ideas to reduce environmental impacts for future subsonic passenger transports. NASA is currently improving and building new analysis capabilities to analyze advanced concepts. To test some of these new capabilities, a transonic truss braced wing configuration was used as a test case. This paper examines the effects due to changes in the design cruise speed and other tradeoffs in the design space. The analysis was baselined to the Boeing SUGAR High truss braced wing concept. An optimization was run at five different design cruise Mach numbers. These designs are compared to provide an initial assessment space and the parameters that should be considered when selecting a design cruise speed. A discussion of the design drivers is also included. The results show that the wing weight in the current analysis has more influence on the takeoff gross weight than expected. This effect caused lower than expected wing sweep angle values for higher cruise speed designs.
Kilgore, R. A.
Experience with the Langley 0.3 meter transonic cryogenic tunnel, which is fan driven, indicated that such a tunnel presents no unusual design difficulties and is simple to operate. Purging, cooldown, and warmup times were acceptable and were predicted with good accuracy. Cooling with liquid nitrogen was practical over a wide range of operating conditions at power levels required for transonic testing, and good temperature distributions were obtained by using a simple liquid nitrogen injection system. To take full advantage of the unique Reynolds number capabilities of the 0.3 meter transonic tunnel, it was designed to accommodate test sections other than the original, octagonal, three dimensional test section. A 20- by 60-cm two dimensional test section was recently installed and is being calibrated. A two dimensional test section with self-streamlining walls and a test section incorporating a magnetic suspension and balance system are being considered.
Vijayaraghavan, D.; Keith, T. G., Jr.; Brewe, D. E.
An analogy between the mathematical modeling of transonic potential flow and the flow in a cavitating bearing is described. Based on the similarities, characteristics of the cavitated region and jump conditions across the film reformation and rupture fronts are developed using the method of weak solutions. The mathematical analogy is extended by utilizing a few computational concepts of transonic flow to numerically model the cavitating bearing. Methods of shock fitting and shock capturing are discussed. Various procedures used in transonic flow computations are adapted to bearing cavitation applications, for example, type differencing, grid transformation, an approximate factorization technique, and Newton's iteration method. These concepts have proved to be successful and have vastly improved the efficiency of numerical modeling of cavitated bearings.
Gallman, J. W.; Batina, J. T.; Yang, T. Y.
An automated flutter boundary tracking procedure for the efficient calculation of transonic flutter boundaries is presented. The procedure uses aeroelastic responses to march along the boundary by taking steps in speed and Mach number, thereby reducing the number of response calculations previously required to determine a transonic flutter boundary. Flutter boundary results are presented for a typical airfoil section oscillating with pitch and plunge degrees of freedom. These transonic flutter boundaries are in good agreement with exact boundaries calculated using the conventional time-marching method. The tracking procedure is extended to include static aeroelastic twist as a simulation of the static deformation of a wing and contains all of the essential features that are required to apply it to practical three-dimensional cases. The procedure is also applied to flutter boundaries as a function of structural parameters.
Rivers, S. M. B.; Wahls, R. A.; Owens, L. R.
A computational study focused on leading-edge radius effects and associated Reynolds number sensitivity for a High Speed Civil Transport configuration at transonic conditions was conducted as part of NASA's High Speed Research Program. The primary purposes were to assess the capabilities of computational fluid dynamics to predict Reynolds number effects for a range of leading-edge radius distributions on a second-generation supersonic transport configuration, and to evaluate the potential performance benefits of each at the transonic cruise condition. Five leading-edge radius distributions are described, and the potential performance benefit including the Reynolds number sensitivity for each is presented. Computational results for two leading-edge radius distributions are compared with experimental results acquired in the National Transonic Facility over a broad Reynolds number range.
Anders, J. B.; Anderson, W. K.; Murthy, A. V.
The use of a high molecular weight test gas to increase the Reynolds number range of transonic wind tunnels is explored. Modifications to a small transonic wind tunnel are described and the real gas properties of the example heavy gas (sulfur hexafluoride) are discussed. Sulfur hexafluoride is shown to increase the test Reynolds number by a factor of more than 2 over air at the same Mach number. Experimental and computational pressure distributions on an advanced supercritical airfoil configuration at Mach 0.7 in both sulfur hexafluoride and nitrogen are presented. Transonic similarity theory is shown to be partially successful in transforming the heavy gas results to equivalent nitrogen (air) results, provided the correct definition of gamma is used.
Deere, Karen A.; Luckring, James M.; McMillin, S. Naomi; Flamm, Jeffrey D.; Roman, Dino
A computational study was performed for a Hybrid Wing Body configuration that was focused at transonic cruise performance conditions. In the absence of experimental data, two fully independent computational fluid dynamics analyses were conducted to add confidence to the estimated transonic performance predictions. The primary analysis was performed by Boeing with the structured overset-mesh code OVERFLOW. The secondary analysis was performed by NASA Langley Research Center with the unstructured-mesh code USM3D. Both analyses were performed at full-scale flight conditions and included three configurations customary to drag buildup and interference analysis: a powered complete configuration, the configuration with the nacelle/pylon removed, and the powered nacelle in isolation. The results in this paper are focused primarily on transonic performance up to cruise and through drag rise. Comparisons between the CFD results were very good despite some minor geometric differences in the two analyses.
Harvey, W. D.; Stainback, P. C.; Owen, F. K.
Wind tunnel testing of low drag airfoils and basic transition studies at transonic speeds are designed to provide high quality aerodynamic data at high Reynolds numbers. This requires that the flow quality in facilities used for such research be excellent. To obtain a better understanding of the characteristics of facility disturbances and identification of their sources for possible facility modification, detailed flow quality measurements were made in two prospective NASA wind tunnels. Experimental results are presented of an extensive and systematic flow quality study of the settling chamber, test section, and diffuser in the Langley 8 foot transonic pressure tunnel and the Ames 12 foot pressure wind tunnel. Results indicate that the free stream velocity and pressure fluctuation levels in both facilities are low at subsonic speeds and are so high as to make it difficult to conduct meaningful boundary layer control and transition studies at transonic speeds.
Curtin, M. M.; Bogue, D. R.; Om, D.; Rivers, S. M. B.; Pendergraft, O. C., Jr.; Wahls, R. A.
This paper discusses Reynolds number scaling for aerodynamic parameters including force and wing pressure measurements. A full-span model of the Boeing 777 configuration was tested at transonic conditions in the National Transonic Facility (NTF) at Reynolds numbers (based on mean aerodynamic chord) from 3.0 to 40.0 million. Data was obtained for a tail-off configuration both with and without wing vortex generators and flap support fairings. The effects of aeroelastics were separated from Reynolds number effects by varying total pressure and temperature independently. Data from the NTF at flight Reynolds number are compared with flight data to establish the wind tunnel/flight correlation. The importance of high Reynolds number testing and the need for developing a process for transonic Reynolds number scaling is discussed. This paper also identifies issues that need to be worked for Boeing Commercial to continue to conduct future high Reynolds number testing in the NTF.
Howlett, James T.
This paper presents a method for calculating viscous effects in two- and three-dimensional unsteady transonic flow fields. An integral boundary-layer method for turbulent viscous flow is coupled with the transonic small-disturbance potential equation in a quasi-steady manner. The viscous effects are modeled with Green's lag-entrainment equations for attached flow and an inverse boundary-layer method for flows that involve mild separation. The boundary-layer method is used stripwise to approximate three-dimensional effects. Applications are given for two-dimensional airfoils, aileron buzz, and a wing planform. Comparisons with inviscid calculations, other viscous calculation methods, and experimental data are presented. The results demonstrate that the present technique can economically and accurately calculate unsteady transonic flow fields that have viscous-inviscid interactions with mild flow separation.
Cunningham, A. M., Jr.
A study was conducted to investigate the feasibility of using combined subsonic and supersonic linear theory as a means for solving unsteady transonic flow problems in an economical and yet realistic manner. With some modification, existing linear theory methods are combined into a single program and a simple algorithm is derived for determining interference between lifting surface elements of different Mach number. The method is applied to a wide variety of problems for which measured unsteady pressure distributions and Mach number distributions are available. By comparing theory and experiment, the transonic method solutions show a significant improvement over uniform flow solutions. It is concluded that with these refinements the method will provide a means for performing realistic transonic flutter and dynamic response analyses at costs which are compatible with current linear theory based solutions.
Reed, W. H., III
A pitch/plunge flutter model suspension system and associated two-dimensional MBB-A3 airfoil models is described. The system is designed for installation in the Langley 6-by-19-inch and 6-by-18-inch transonic blowdown wind tunnels to enable systematic study of the transonic flutter characteristics and static pressure distributions of supercritical airfoils at transonic Mach numbers. A compound spring suspension concept is introduced which simultaneously meets requirements for low plunge-mode stiffness, lightweight suspended model, and large steady lift due to angle of attack without the need for excessive static deflections of the plunge spring. The system features variable pitch and plunge frequencies, changeable airfoil rotation axes, and a self aligning control system to maintain a constant mean position of the model with changing airload.
Yaros, S. F.; Carlson, J. R.; Chandrasekaran, B.
An effort has been undertaken at the NASA Langley Research Center to assess the capabilities of available computational methods for use in propulsion integration design studies of transonic transport aircraft, particularly of pylon/nacelle combinations which exhibit essentially no interference drag. The three computer codes selected represent state-of-the-art computational methods for analyzing complex configurations at subsonic and transonic flight conditions. These are: EULER, a finitie volume solution of the Euler equation; VSAERO, a panel solution of the Laplace equation; and PPW, a finite difference solution of the small disturbance transonic equations. In general, all three codes have certain capabilities that allow them to be of some value in predicting the flows about transport configurations, but all have limitations. Until more accurate methods are available, careful application and interpretation of the results of these codes are needed.
Yaros, Steven F.; Carlson, John R.; Chandrasekaran, Balasubramanyan
An effort has been undertaken at the NASA Langley Research Center to assess the capabilities of available computational methods for use in propulsion integration design studies of transonic transport aircraft, particularly of pylon/nacelle combinations which exhibit essentially no interference drag. The three computer codes selected represent state-of-the-art computational methods for analyzing complex configurations at subsonic and transonic flight conditions. These are: EULER, a finite volume solution of the Euler equation; VSAERO, a panel solution of the Laplace equation; and PPW, a finite difference solution of the small disturbance transonic equations. In general, all three codes have certain capabilities that allow them to be of some value in predicting the flows about transport configurations, but all have limitations. Until more accurate methods are available, careful application and interpretation of the results of these codes are needed.
Seidel, D. A.; Sandford, M. C.; Eckstrom, C. V.
Transonic steady and unsteady aerodynamic data were measured on a large elastic wing in the NASA Langley Transonic Dynamics Tunnel. The wing had a supercritical airfoil shape and a leading-edge sweepback of 28.8 deg. The wing was heavily instrumented to measure both static and dynamic pressures and deflections. A hydraulically driven outboard control surface was oscillated to generate unsteady airloads on the wing. Representative results from the wind tunnel tests are presented and discussed, and the unexpected occurrence of an unusual dynamic wing instability, which was sensitive to angle of attack, is reported.
Seidel, D. A.; Bennett, R. M.; Whitlow, W., Jr.
A pulse-transfer function technique for calculating unsteady aerodynamic forces for a wide range of reduced frequencies is implemented in a finite difference program solving the complete unsteady transonic small perturbation equation. Forces are calculated for a two-dimensional linear flat plate case utilizing the default grids from several currently used finite difference programs. The forces are compared to exact theoretical values and grid generated boundary and internal reflections are demonstrated. Grids designed to alleviate the reflections are presented and forces for a 6% thick parabolic arc airfoil are calculated to investigate non-linear transonic effects.
Mercer, Charles E.; Berrier, Bobby L.; Capone, Francis J.; Grayston, Alan M.; Sherman, C. D.
The equations used by the 16 foot transonic tunnel in the data reduction programs are presented in eight modules. Each module consists of equations necessary to achieve a specific purpose. These modules are categorized in the following groups: tunnel parameters; jet exhaust measurements; skin friction drag; balance loads and model attitudes calculations; internal drag (or exit-flow distributions); pressure coefficients and integrated forces; thrust removal options; and turboprop options. This document is a companion document to NASA TM-83186, A User's Guide to the Langley 16 Foot Transonic Tunnel, August 1981.
Yacoby, Eyal; Sadot, Oren; Barmashenko, Boris D.; Rosenwaks, Salman
Three-dimensional computational fluid dynamics (3D CFD) modeling of subsonic (Mach number M ~ 0.2) and transonic (M ~ 0.9) diode pumped alkali lasers (DPALs), taking into account fluid dynamics and kinetic processes in the lasing medium is reported. The performance of these lasers is compared with that of supersonic (M ~ 2.7 for Cs and M ~ 2.4 for K) DPALs. The motivation for this study stems from the fact that subsonic and transonic DPALs require much simpler hardware than supersonic ones where supersonic nozzle, diffuser and high power mechanical pump (due to a drop in the gas total pressure in the nozzle) are required for continuous closed cycle operation. For Cs DPALs with 5 x 5 cm2 flow cross section pumped by large cross section (5 x 2 cm2) beam the maximum achievable power of supersonic devices is higher than that of the transonic and subsonic devices by only ~ 3% and ~ 10%, respectively. Thus in this case the supersonic operation mode has no substantial advantage over the transonic one. The main processes limiting the power of Cs supersonic DPALs are saturation of the D2 transition and large ~ 60% losses of alkali atoms due to ionization, whereas the influence of gas heating is negligible. For K transonic DPALs both the gas heating and ionization effects are shown to be unimportant. The maximum values of the power are higher than those in Cs transonic laser by ~ 11%. The power achieved in the supersonic and transonic K DPAL is higher than for the subsonic version, with the same resonator and K density at the inlet, by ~ 84% and ~ 27%, respectively, showing a considerable advantaged of the supersonic device over the transonic one. For pumping by rectangular beams of the same (5 x 2 cm2) cross section, comparison between end-pumping - where the laser beam and pump beam both propagate at along the same axis, and transverse-pumping - where they propagate perpendicularly to each other, shows that the output power and optical-to-optical efficiency are not
Baker, A. J.; Fox, C. H., Jr.
Routine determination of inviscid subsonic flow fields about wing-body-tail configurations employing a Green's function approach for numerical solution of the perturbation velocity potential equation is successfully extended into the high subsonic subcritical flow regime and into the shock-free supersonic flow regime. A modified Green's function formulation, valid throughout a range of Mach numbers including transonic, that takes an explicit accounting of the intrinsic nonlinearity in the parent governing partial differential equations is developed. Some considerations pertinent to flow field predictions in the transonic flow regime are discussed.
Gaffney, R. L., Jr.; Hassan, H. A.; Salas, M. D.
A finite volume formulation of the Euler equations using Cartesian grids is used to calculate the transonic flow over multielement airfoils and to use the resulting solutions to assess wall interference effects in wind tunnels. Available methods and recommendations for evaluating such effects, which are based on shifts in Mach number and angle of attack, are examined and the results are compared with measurements using the flapped supercritical SKF 1.1 airfoil. Based on the calculations, it is concluded that shifts in Mach number and angle of attack cannot by themselves account for viscous and wall effects on multielement airfoils at transonic speeds.
Roberts, W. B.; Vanrintel, H. L.; Rizvi, G.
A simulated transonic rotor channel model was examined experimentally to verify the flow physics of internal area ruling. Pressure measurements were performed in the high speed wind tunnel at transonic speeds with Mach 1.5 and Mach 2 nozzle blocks to get an indication of the approximate shock losses. The results showed a reduction in losses due to internal area ruling with the Mach 1.5 nozzle blocks. The reduction in total loss coefficient was of the order of 17 percent for a high blockage model and 7 percent for a cut-down model.
Lindsey, Walter F; Landrum, Emma Jean
Schlieren photographs have been compiled of the two-dimensional flow at transonic speeds past 37 airfoils. These airfoils have variously shaped profiles, and some are related in thickness and camber. The data for these airfoils were analyzed to provide basic information on the flow changes involved and to determine factors affecting transonic-flow attachment, which is a transition from separated to unseparated flow at the leading edges of two-dimensional airfoils at fixed angles as the subsonic Mach number is increased.
Guo, Tongqing; Lu, Zhiliang; Wu, Yongjian
Gridgen is employed for static multi-block grid generation. A rapid deforming technique is employed for dynamic grids. Flutters are numerically analyzed in the time domain with a coupled solution of unsteady Euler equations and structural equations of motion. Based on variable stiffness method of transonic flutter analysis, a time-domain method of transonic flutter analysis with multi-directional coupled vibrations is develpoed. For completeness, flutter characteristics of a wing model with winglets and an aircraft model with external stores are numerically analyzed.
Holst, T. L.; Ballhaus, W. F.
Implicit approximate factorization techniques (AF) were investigated for the solution of matrix equations resulting from finite difference approximations to the full potential equation in conservation form. For transonic flows, an artificial viscosity, required to maintain stability in supersonic regions, was introduced by an upwind bias of the density. Two implicit AF procedures are presented and their convergence performance is compared with that of the standard transonic solution procedure, successive line overrelaxation (SLOR). Subcritical and supercritical test cases are considered. The results indicate that the AF schemes are substantially faster than SLOR.
Aerodynamic design optimization of transonic fan blades is a highly challenging problem due to the complexity of flow field inside the fan, the conflicting design requirements and the high-dimensional design space. In order to address all these challenges, an aerodynamic design optimization method is developed in this study. This method automates the design process by integrating a geometrical parameterization method, a CFD solver and numerical optimization methods that can be applied to both single and multi-point optimization design problems. A multi-level blade parameterization is employed to modify the blade geometry. Numerical analyses are performed by solving 3D RANS equations combined with SST turbulence model. Genetic algorithms and hybrid optimization methods are applied to solve the optimization problem. In order to verify the effectiveness and feasibility of the optimization method, a singlepoint optimization problem aiming to maximize design efficiency is formulated and applied to redesign a test case. However, transonic fan blade design is inherently a multi-faceted problem that deals with several objectives such as efficiency, stall margin, and choke margin. The proposed multi-point optimization method in the current study is formulated as a bi-objective problem to maximize design and near-stall efficiencies while maintaining the required design pressure ratio. Enhancing these objectives significantly deteriorate the choke margin, specifically at high rotational speeds. Therefore, another constraint is embedded in the optimization problem in order to prevent the reduction of choke margin at high speeds. Since capturing stall inception is numerically very expensive, stall margin has not been considered as an objective in the problem statement. However, improving near-stall efficiency results in a better performance at stall condition, which could enhance the stall margin. An investigation is therefore performed on the Pareto-optimal solutions to
Nagabushana, K. A.; Ash, Robert L.
A different approach for calibrating hot-wires, which simplifies the calibration procedure and reduces the tunnel run-time by an order of magnitude was sought. In general, it is accepted that the directly measurable quantities in any flow are velocity, density, and total temperature. Very few facilities have the capability of varying the total temperature over an adequate range. However, if the overheat temperature parameter, a(sub w), is used to calibrate the hot-wire then the directly measurable quantity, voltage, will be a function of the flow variables and the overheat parameter i.e., E = f(u,p,a(sub w), T(sub w)) where a(sub w) will contain the needed total temperature information. In this report, various methods of evaluating sensitivities with different dependent and independent variables to calibrate a 3-Wire hot-wire probe using a constant temperature anemometer (CTA) in subsonic/transonic flow regimes is presented. The advantage of using a(sub w) as the independent variable instead of total temperature, t(sub o), or overheat temperature parameter, tau, is that while running a calibration test it is not necessary to know the recovery factor, the coefficients in a wire resistance to temperature relationship for a given probe. It was deduced that the method employing the relationship E = f (u,p,a(sub w)) should result in the most accurate calibration of hot wire probes. Any other method would require additional measurements. Also this method will allow calibration and determination of accurate temperature fluctuation information even in atmospheric wind tunnels where there is no ability to obtain any temperature sensitivity information at present. This technique greatly simplifies the calibration process for hot-wires, provides the required calibration information needed in obtaining temperature fluctuations, and reduces both the tunnel run-time and the test matrix required to calibrate hotwires. Some of the results using the above techniques are presented
Sekula, Martin K.; Piatak, David J.; Rausch, Russ D.
A transonic wind tunnel test of the Ares I-X Rigid Buffet Model (RBM) identified a Mach number regime where unusually large buffet loads are present. A subsequent investigation identified the cause of these loads to be an alternating flow phenomenon at the Crew Module-Service Module junction. The conical design of the Ares I-X Crew Module and the cylindrical design of the Service Module exposes the vehicle to unsteady pressure loads due to the sudden transition between a subsonic separated and a supersonic attached flow about the cone-cylinder junction as the local flow randomly fluctuates back and forth between the two flow states. These fluctuations produce a square-wave like pattern in the pressure time histories resulting in large amplitude, impulsive buffet loads. Subsequent testing of the Ares I RBM found much lower buffet loads since the evolved Ares I design includes an ogive fairing that covers the Crew Module-Service Module junction, thereby making the vehicle less susceptible to the onset of alternating flow. An analysis of the alternating flow separation and attachment phenomenon indicates that the phenomenon is most severe at low angles of attack and exacerbated by the presence of vehicle protuberances. A launch vehicle may experience either a single or, at most, a few impulsive loads since it is constantly accelerating during ascent rather than dwelling at constant flow conditions in a wind tunnel. A comparison of a windtunnel- test-data-derived impulsive load to flight-test-data-derived load indicates a significant over-prediction in the magnitude and duration of the buffet load. I. Introduction One
The Prandtl-Busemann small-perturbation method is utilized to obtain the flow of a compressible fluid past an infinitely long wave-shaped wall. When the essential assumption for transonic flow (that all Mach numbers in the region of flow are nearly unity) is introduced, the expression for the velocity potential takes the form of a power series in the transonic similarity parameter. On the basis of this form of the solution, an attempt is made to solve the nonlinear differential equation for transonic flow past the wavy wall. The analysis utilized exhibits clearly the difficulties inherent in nonlinear-flow problems.
Snow, W. L.; Burner, A. W.; Goad, W. K.
Viewing problems associated with the measurement of model deformation in cryogenic wind tunnels are discussed. Tests were conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel to assess viewing capabilities thru the flow field. The effects of condensation and turbulent boundary layers are discussed and a modelling procedure for image degradation is described.
Chang, J. F.; Lan, C. Edward
The use of a transonic airfoil code for analysis, inverse design, and direct optimization of an airfoil immersed in propfan slipstream is described. A summary of the theoretical method, program capabilities, input format, output variables, and program execution are described. Input data of sample test cases and the corresponding output are given.
Rizzetta, D. P.
Modifications of the code LTRAN2, developed by Ballhaus and Goorjian, which account for viscous effects in the computation of planar unsteady transonic flows are presented. Two models are considered and their theoretical development and numerical implementation is discussed. Computational examples employing both models are compared with inviscid solutions and with experimental data. Use of the modified code is described.
Wahls, Richard A.; Owens, Lewis R., Jr.; Londenberg, W. Kelly
Experience with afterbody closure effects and accompanying test techniques issues on a High Speed Civil Transport (HSCT)-class configuration is described. An experimental data base has been developed which includes force, moment, and surface pressure data for the High Speed Research (HSR) Reference H configuration with a closed afterbody at subsonic and transonic speeds, and with a cylindrical afterbody at transonic and supersonic speeds. A supporting computational study has been performed using the USM3D unstructured Euler solver for the purposes of computational fluid dynamics (CFD) method assessment and model support system interference assessment with a focus on lower blade mount effects on longitudinal data at transonic speeds. Test technique issues related to a lower blade sting mount strategy are described based on experience in the National Transonic Facility (NTF). The assessment and application of the USM3D code to the afterbody/sting interference problem is discussed. Finally, status and plans to address critical test technique issues and for continuation of the computational study are presented.
Newman, James C., III; Baysal, Oktay
Transonic Euler calculations about a complex multicomponent configuration are presented. The 3D Euler equations are solved utilizing an upwind-biased, alternating direction implicit, approximately factored, multigrid algorithm. Computational results are compared to experimental data of the finned store in a carriage position.
Batina, John T.; Seidel, David A.; Bennett, Robert M.; Cunningham, Herbert J.; Bland, Samuel R.
A transonic unsteady aerodynamic and aeroelastic code called CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) was developed for application to realistic aircraft configurations. It permits the calculation of steady and unsteady flows about complete aircraft configurations for aeroelastic analysis of the flutter critical transonic speed range. The CAP-TSD code uses a time accurate approximate factorization algorithm for solution of the unsteady transonic small disturbance potential equation. An overview is given of the CAP-TSD code development effort along with recent algorithm modifications which are listed and discussed. Calculations are presented for several configurations including the General Dynamics 1/9th scale F-16C aircraft model to evaluate the algorithm and hence the reliability of the CAP-TSD code in general. Calculations are also presented for a flutter analysis of a 45 deg sweptback wing which agree well with the experimental data. Descriptions are presented of the CAP-TSD code and algorithm details along with results and comparisons which demonstrate the stability, accuracy, efficiency, and utility of CAP-TSD.
A finite element method for the computation of the transonic flow with shocks past airfoils is presented using the artificial viscosity concept for the local supersonic regime. Generally, the classic element types do not meet the accuracy requirements of advanced numerical aerodynamics requiring special attention to the choice of an appropriate element. A series of computed pressure distributions exhibits the usefulness of the method.
The use of a fast elliptic solver in combination with relaxation is presented as an effective way to accelerate the convergence of transonic flow calculations, particularly when a marching scheme can be used to treat the supersonic zone in the relaxation process.
Heaslet, Max A; Spreiter, John R
The present paper re-examines the derivation of the integral equations for transonic flow around slender wings and bodies of revolution, giving special attention to conditions resulting from the presence of shock waves and to the reduction of the relations to the special forms necessary for the discussion of sonic flow, that is, flow at free-stream Mach number 1.
Mead, H. R.; Melnik, R. E.
A user's manual which describes the operation of the computer program, GRUMFOIL is presented. The program computes the viscous transonic flow over two dimensional airfoils using a boundary layer type viscid-inviscid interaction approach. The inviscid solution is obtained by a multigrid method for the full potential equation. The boundary layer solution is based on integral entrainment methods.
Corson, B. W., Jr.; Runckel, J. F.; Igoe, W. B.
The Langley 16-foot transonic tunnel with test section air removal (plenum suction) was calibrated to a Mach number of 1.3. The results of the calibration, including the effects of slot shape modifications, test section wall divergence, and water vapor condensation, are presented. A complete description of the wind tunnel and its auxiliary equipment is included.
indicators of symmetry since the wings were unbanked within the limits of tolerances and flow angularity. Longitudinal, spanwise, and vertical... unbanked wings at subsonic and transonic speeds from low to high angles of attack. The wing planforms varied in aspect ratio and taper ratio with
UNCLASSIFIED UNCLASSIFIED Metrology Measurements of the DSTO Transonic Wind Tunnel Store Support Arm Adam Blandford, John Clayton...provide quality assurance for test clients. This document details metrology measurements that were conducted during February and March 2013 on the store...Australia 2013 AR-015-818 December 2013 APPROVED FOR PUBLIC RELEASE UNCLASSIFIED UNCLASSIFIED Metrology Measurements of the DSTO
The operational characteristics and equipment associated with the Langley 16-foot transonic tunnel complex which is located in buildings 1146 and 1234 at the Langley Research Center are described in detail. This complex consists of the 16-foot transonic wind tunnel, the static test facility, and the 16- by 24-inch water tunnel research facilities. The 16-foot transonic tunnel is a single-return atmospheric wind tunnel with a 15.5 foot diameter test section and a Mach number capability from 0.20 to 1.30. The emphasis for research conducted in this research complex is on the integration of the propulsion system into advanced aircraft concepts. In the past, the primary focus has been on the integration of nozzles and empennage into the afterbody of fighter aircraft. During the last several years this experimental research has been expanded to include developing the fundamental data base necessary to verify new theoretical concepts, inlet integration into fighter aircraft, nozzle integration for supersonic and hypersonic transports, nacelle/pylon/wing integration for subsonic transport configurations, and the study of vortical flows (in the 16- by 24-inch water tunnel). The purpose here is to provide a comprehensive description of the operational characteristics of the research facilities of the 16-foot transonic tunnel complex and their associated systems and equipments.
Chintsun, H.; Pi, W. S.
A flight test and data processing program utilizing a Northrop F-5A aircraft instrumented to acquire buffet pressures and response data during transonic maneuvers is discussed. The data are presented in real-time format followed by spectral and statistical analyses. Also covered is a comparison of the aircraft response data with computed responses based on the measured buffet pressures.
Schaefer, John; West, Thomas; Hosder, Serhat; Rumsey, Christopher; Carlson, Jan-Renee; Kleb, William
The goal of this work was to quantify the uncertainty and sensitivity of commonly used turbulence models in Reynolds-Averaged Navier-Stokes codes due to uncertainty in the values of closure coefficients for transonic, wall-bounded flows and to rank the contribution of each coefficient to uncertainty in various output flow quantities of interest. Specifically, uncertainty quantification of turbulence model closure coefficients was performed for transonic flow over an axisymmetric bump at zero degrees angle of attack and the RAE 2822 transonic airfoil at a lift coefficient of 0.744. Three turbulence models were considered: the Spalart-Allmaras Model, Wilcox (2006) k-w Model, and the Menter Shear-Stress Trans- port Model. The FUN3D code developed by NASA Langley Research Center was used as the flow solver. The uncertainty quantification analysis employed stochastic expansions based on non-intrusive polynomial chaos as an efficient means of uncertainty propagation. Several integrated and point-quantities are considered as uncertain outputs for both CFD problems. All closure coefficients were treated as epistemic uncertain variables represented with intervals. Sobol indices were used to rank the relative contributions of each closure coefficient to the total uncertainty in the output quantities of interest. This study identified a number of closure coefficients for each turbulence model for which more information will reduce the amount of uncertainty in the output significantly for transonic, wall-bounded flows.
Luckring, J. M.
A review of National Transonic Facility (NTF) investigations for high-performance military aerodynamics has been completed. The review spans the entire operational period of the tunnel, and includes configurations ranging from full aircraft to basic research geometries. The intent for this document is to establish a comprehensive summary of these experiments with selected technical results
Wai, J. C.; Sun, C. C.; Yoshihara, H.
A method using a transonic small disturbance code with successive line over-relaxation is described for treating wing/fuselage configurations with a nacelle/pylon/powered jet. Examples illustrating its use for the NASA transport research model are given. Reasonable test/theory comparisons were obtained.
Hafez, M. M.; Wellford, L. C.; Merkle, C. L.; Murman, E. M.
Studies on applications of the finite element approach to transonic flow calculations are reported. Different discretization techniques of the differential equations and boundary conditions are compared. Finite element analogs of Murman's mixed type finite difference operators for small disturbance formulations were constructed and the time dependent approach (using finite differences in time and finite elements in space) was examined.
character of the hodograph equations. one-dimensional, transonic flow in a laval nozzle in the vicinity of the throat was obtained in the For such...state of development, the hodograph method has 2 limitations. It can only be applied to the design a (3) problem in plane flow and cannot be used in
Griffin, S. A.; Madsen, A. P.; Mcclain, A. A.
The feasibility of designing advanced technology, highly maneuverable, fighter aircraft models to achieve full scale Reynolds number in the National Transonic Facility (NTF) is examined. Each of the selected configurations are tested for aeroelastic effects through the use of force and pressure data. A review of materials and material processes is also included.
Igarashi, Asuka; Mori, Masao; Nitta, Shin-Ya
We study fundamental properties of transonic galactic outflows in the gravitational potential of a cold dark matter halo (DMH) with a central super-massive black hole (SMBH) assuming a polytropic, steady and spherically symmetric state. We have classified the transonic solutions with respect to their topology in the phase space. As a result, we have found two types of transonic solutions characterized by a magnitude relationship between the gravity of DMH and that of SMBH. These two types of solutions have different loci of the transonic points; one transonic point is formed at a central region (< 0.01kpc) and another is at a distant region (> 100kpc). Also, mass fluxes and outflow velocities are different between the two solutions. These two transonic solutions may play different roles on the star formation history of galaxies and the metal contamination of intergalactic space. Furthermore, we have applied our model to the Sombrero galaxy. In this galaxy, the wide-spread hot gas is detected as an apparent trace of galactic outflows while the star-formation rate is disproportionately low, and the observed gas density distribution is quite similar to the hydrostatic state (Li et al. 2011). To solve this discrepancy, we propose a slowly accelerating outflow in which the transonic point forms in a distant region (~ 120 kpc) and the subsonic region spreads across the stellar distribution. In the subsonic region, the gas density distribution is similar to that of the hydrostatic state. Our model predicts the possibility of the slowly accelerating outflow in the Sombrero galaxy. Igarashi et al. 2014 used the isothermal model and well reproduced the observed gas density distribution, but the estimated mass flux (1.8M ⊙/yr) is lager than the mass of the gas supplied by stars (0.3-0.4M ⊙/yr). Then, we expect that the polytropic model may reproduce the observational mass of the supplied gas (Igarashi et al. 2015). Such slowly accelerating outflows should be distinguished
Pak, Chan-gi; Li, Wesley W.
Supporting the Aeronautics Research Mission Directorate guidelines, the National Aeronautics and Space Administration [NASA] Dryden Flight Research Center is developing a multidisciplinary design, analysis, and optimization [MDAO] tool. This tool will leverage existing tools and practices, and allow the easy integration and adoption of new state-of-the-art software. Today s modern aircraft designs in transonic speed are a challenging task due to the computation time required for the unsteady aeroelastic analysis using a Computational Fluid Dynamics [CFD] code. Design approaches in this speed regime are mainly based on the manual trial and error. Because of the time required for unsteady CFD computations in time-domain, this will considerably slow down the whole design process. These analyses are usually performed repeatedly to optimize the final design. As a result, there is considerable motivation to be able to perform aeroelastic calculations more quickly and inexpensively. This paper will describe the development of unsteady transonic aeroelastic design methodology for design optimization using reduced modeling method and unsteady aerodynamic approximation. The method requires the unsteady transonic aerodynamics be represented in the frequency or Laplace domain. Dynamically linear assumption is used for creating Aerodynamic Influence Coefficient [AIC] matrices in transonic speed regime. Unsteady CFD computations are needed for the important columns of an AIC matrix which corresponded to the primary modes for the flutter. Order reduction techniques, such as Guyan reduction and improved reduction system, are used to reduce the size of problem transonic flutter can be found by the classic methods, such as Rational function approximation, p-k, p, root-locus etc. Such a methodology could be incorporated into MDAO tool for design optimization at a reasonable computational cost. The proposed technique is verified using the Aerostructures Test Wing 2 actually designed
Chokani, Ndaona; Milholen, William E., II
A semi-span testing technique has been proposed for the NASA Langley Research Center's National Transonic Facility (NTF). Semi-span testing has several advantages including (1) larger model size, giving increased Reynolds number capability; (2) improved model fidelity, allowing ease of flap and slat positioning which ultimately improves data quality; and (3) reduced construction costs compared with a full-span model. In addition, the increased model size inherently allows for increased model strength, reducing aeroelastic effects at the high dynamic pressure levels necessary to simulate flight Reynolds numbers. The Energy Efficient Transport (EET) full-span model has been modified to become the EET semi-span model. The full-span EET model was tested extensively at both NASA LRC and NASA Ames Research Center. The available full-span data will be useful in validating the semi-span test strategy in the NTF. In spite of the advantages discussed above, the use of a semi-span model does introduce additional challenges which must be addressed in the testing procedure. To minimize the influence of the sidewall boundary layer on the flow over the semi-span model, the model must be off-set from the sidewall. The objective is to remove the semi-span model from the sidewall boundary layer by use of a stand-off geometry. When this is done however, the symmetry along the centerline of the full-span model is lost when the semi-span model is mounted on the wind tunnel sidewall. In addition, the large semi-span model will impose a significant pressure loading on the sidewall boundary layer, which may cause separation. Even under flow conditions where the sidewall boundary layer remains attached, the sidewall boundary layer may adversely effect the flow over the semi-span model. Also, the increased model size and sidewall mounting requires a modified wall correction strategy. With these issues in mind, the semi-span model has been well instrumented with surface pressure taps to
Rivers, Melissa B.; Quest, Jurgen; Rudnik, Ralf
Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility, the NASA Ames 11-ft wind tunnel, and the European Transonic Wind Tunnel. In the NASA Ames 11-ft wind tunnel, data have been obtained at only a chord Reynolds number of 5 million for a wing/body/tail = 0 degree incidence configuration. Data have been obtained at chord Reynolds numbers of 5, 19.8 and 30 million for the same configuration in the National Transonic Facility and in the European Transonic Facility. Force and moment, surface pressure, wing bending and twist, and surface flow visualization data were obtained in all three facilities but only the force and moment, surface pressure and wing bending and twist data are presented herein.
Rivers, Melissa; Quest, Juergen; Rudnik, Ralf
Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility, the NASA Ames 11-ft wind tunnel, and the European Transonic Wind Tunnel. In the NASA Ames 11-ft wind tunnel, data have been obtained at only a chord Reynolds number of 5 million for a wing/body/tail = 0 degree incidence configuration. Data have been obtained at chord Reynolds numbers of 5, 19.8 and 30 million for the same configuration in the National Transonic Facility and in the European Transonic Facility. Force and moment, surface pressure, wing bending and twist, and surface flow visualization data were obtained in all three facilities but only the force and moment and surface pressure data are presented herein.
Bennett, Robert M.; Bland, Samuel R.; Batina, John T.; Gibbons, Michael D.; Mabey, Dennis G.
A transonic unsteady aerodynamic and aeroelasticity code has been developed for application to realistic aircraft configurations. The new code is called CAP-TSD which is an acronym for Computational Aeroelasticity Program - Transonic Small Disturbance. The CAP-TSD code uses a time-accurate approximate factorization algorithm for solution of the unsteady transonic small-disturbance equation that is efficient for solution of steady and unsteady transonic flow problems including supersonic freestream flows. The new code can treat complete aircraft geometries with multiple lifting surfaces and bodies. Applications to wings in supersonic freestream flow are presented. Comparisons with selected exact solutions from linear theory are presented showing generally favorable results. Calculations for both steady and oscillatory cases for the F-5 and RAE tailplane models are compared with experimental data and also show good overall agreement. Selected steady calculations are further compared with a steady flow Euler code.
release; distribution is unlimited. Understanding and Migitating Vortex-Dominated, Tip-Leakage and End- Wall Losses in a Transonic Splittered Rotor Stage...Circle Monterey, CA 93943 -5000 ABSTRACT Understanding and Migitating Vortex-Dominated, Tip-Leakage and End- Wall Losses in a Transonic Splittered Rotor...mitigate tip leakage and end wall flows. Since computational analysis has matured to the level where design optimization can be performed with such
Bennett, R. M.; Wynne, E. C.; Mabey, D. G.
Pressure data measured by the British Royal Aircraft Establishment for the AGARD SMP tailplane are compared with results calculated using the transonic small perturbation code XTRAN3S. A brief description of the analysis is given and a recently developed finite difference grid is described. Results are presented for five steady and nine harmonically oscillating cases near zero angle of attack and for a range of subsonic and transonic Mach numbers.
A dynamic response test performed in a eight foot transonic pressure tunnel is described. The dynamics of the flow process of the wind tunnel at transonic conditions were obtained. Descriptions of the test conditions, instrumentation, presentation of raw data, analysis of data, and finally, based on experimental evidences, an attempt to construct an input output relationship of the flow process from the viewpoints of control engineering are included.
Silva, Walter A.; Ringertz, Ulf; Stenfelt, Gloria; Eller, David; Keller, Donald F.; Chwalowski, Pawel
This paper presents a status report on the collaboration between the Royal Institute of Technology (KTH) in Sweden and the NASA Langley Research Center regarding the design, fabrication, modeling, and testing of a full-span lighter configuration in the Transonic Dynamics Tunnel (TDT). The goal of the test is to acquire transonic limit-cycle- oscillation (LCO) data, including accelerations, strains, and unsteady pressures. Finite element models (FEMs) and aerodynamic models are presented and discussed along with results obtained to date.
The Aerodynamics Division at NASA ARC is participating in the propulsion airframe integration phase of the High Speed Research Program. The two areas of research being pursued include an experimental program and analysis using computational fluid dynamics. The applied Aerodynamics Branch is conducting the experimental program, which will involve a nacelle airframe model that was tested in the Ames 11- by 11-Foot Transonic Wind Tunnel in 1973. This branch will also assess various Euler codes in predicting nacelle airframe interference effects. The goal is to provide the industry with the necessary data and tools to design a high speed civil transport with favorable propulsion airframe interference.
Modern turbofan engines employ a highly loaded fan stage with transonic or low-supersonic velocities in the blade-tip region. The fan blades are often prone to flutter at off-design conditions. Flutter is a highly undesirable and dangerous self-excited mode of blade oscillations that can result in high-cycle fatigue blade failure. The origins of blade flutter are not fully understood yet. The latest view is that the blade oscillations are triggered by high-frequency changes in the extent of the partially separated area on the airfoil suction side. There is a lack of experimental data describing the separated flow characteristics of modern airfoils for transonic fans.
Wolf, S. W. D.
Work was continued with the low speed self streamlining wind tunnel (SSWT) using the NACA 0012-64 airfoil in an effort to explain the discrepancies between the NASA Langley low turbulence pressure tunnel (LTPT) and SSWT results obtained with the airfoil stalled. Conventional wind tunnel corrections were applied to straight wall SSWT airfoil data, to illustrate the inadequacy of standard correction techniques in circumstances of high blockage. Also one SSWT test was re-run at different air speeds to investigate the effects of such changes (perhaps through changes in Reynold's number and freestream turbulence levels) on airfoil data and wall contours. Mechanical design analyses for the transonic self-streamlining wind tunnel (TSWT) were completed by the application of theoretical airfoil flow field data to the elastic beam and streamline analysis. The control system for the transonic facility, which will eventually allow on-line computer operation of the wind tunnel, was outlined.
Silva, Walter A.; Bennett, Robert M.
The CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) code, developed at the NASA-Langley Research Center, is applied to the Active Flexible Wing (AFW) wind-tunnel model for prediction of the model's transonic aeroelastic behavior. Static aeroelastic solutions using CAP-TSD are computed. Dynamic (flutter) analyses are then performed as perturbations about the static aeroelastic deformations of the AFW. The accuracy of the static aeroelastic procedure is investigated by comparing analytical results to those from AFW wind-tunnel experiments. Dynamic results are presented in the form of root loci at different Mach numbers for a heavy gas and for air test mediums. The resultant flutter boundaries for both gases, and the effects of viscous damping and angle of attack on the flutter boundary in air, are also presented.
Lawing, P. L.; Johnson, C. B.
This paper describes test techniques used to obtain data in the 0.3-m Transonic Cryogenic Tunnel. Main sections include steady aerodynamic testing, unsteady aerodynamics, non-intrusive measurements, tunnel performance, and fluid mechanics. Test techniques with adequate previous documentation are briefly presented and referenced, and those for which no documentation yet exists are more thoroughly discussed. Attention is given to the model building and instrumentation technology necessary for testing at high Reynolds numbers in the presence of free transition. It is concluded that the testing techniques thus far demonstrated in the 0.3-m TCT are on par with modern transonic tunnels, and that the specialized techniques needed to exploit the advantages of cryogenic operation can be realized.
Ray, E. J.; Ladson, C. L.; Adcock, J. B.; Lawing, P. L.; Hall, R. M.
The past 6 years of operation with the NASA Langley 0.3 m transonic cryogenic tunnel (TCT) show that there are no insurmountable problems associated with cryogenic testing with gaseous nitrogen at transonic Mach numbers. The fundamentals of the concept were validated both analytically and experimentally and the 0.3 m TCT, with its unique Reynolds number capability, was used for a wide variety of aerodynamic tests. Techniques regarding real-gas effects were developed and cryogenic tunnel conditions can be set and maintained accurately. Cryogenic cooling by injecting liquid nitrogen directly into the tunnel circuit imposes no problems with temperature distribution or dynamic response characteristics. Experience with the 0.3 m TCT, indicates that there is a significant learning process associated with cryogenic, high Reynolds number testing. Many of the questions have already been answered; however, factors such as tunnel control, run logic, economics, instrumentation, and model technology present many new and challenging problems.
Bryanston-Cross, P. J.; Judge, T. R.; Quan, C.; Pugh, G.; Corby, N.
The paper is an extension to earlier PIV work published in Progress in Aerospace Science Vol. 27, pp237-265. 1990. DPIV (Digital Particle Image Velocimetry) measurements have been made at transonic speeds using a new method of both digitally capturing and visualising high speed flow-fields. It has provided a quantitative whole field image of the transonic flow. The area of interest for this test was of the wing/pylon/nacelle/inboard gully region of a 5.7% scale model wing and engine combination. The measurements show velocity and flow angle changes in this region, to a measurement accuracy of 4%, made at an optical stand off distance of 0.5 m.
Weatherill, W. H.; Ehlers, F. E.
Typical section flutter solutions are obtained using generalized air forces calculated with OPTRAN2, a program based on the transonic small perturbation potential equation, which uses potentials from a nonlinear solution to the steady flow to obtain linear unsteady air forces. Flutter results are calculated with a conventional, linear V-g type solution. Results are presented for a NACA 64A010 airfoil oscillating in plunge and pitch, and with a quarter-chord control surface in the Mach number range of 0.80 to 0.90. The flutter boundary shows a significant 'transonic' bucket, together with changes in flutter mode as, with increasing Mach numbers, a shock appears on the airfoil and then moves aft. Single degree-of-freedom instabilities associated with both the pitch and the control surface motions are identified.
Smith, C. W.; Braymen, W. W.; Bhateley, I. C.; Londenberg, W. K.
Numerous computational fluid dynamics (CFD) codes are available that solve any of several variations of the transonic flow equations from small disturbance to full Navier-Stokes. The design philosophy at General Dynamics Fort Worth Division involves use of all these levels of codes, depending on the stage of configuration development. Throughout this process, drag calculation is a central issue. An overview is provided for several transonic codes and representative test-to-theory comparisons for fighter-type configurations are presented. Correlations are shown for lift, drag, pitching moment, and pressure distributions. The future of applied CFD is also discussed, including the important task of code validation. With the progress being made in code development and the continued evolution in computer hardware, the routine application of these codes for increasingly more complex geometries and flow conditions seems apparent.
Dreim, D. R.; Jacobson, S. B.; Britt, R. T.
At high subsonic flight speeds, large flexible aircraft begin to encounter unsteady airloads which are not predicted by most currently available aerodynamic analysis and design methods. With increasing speed and the development of transonic flow and shocks, viscous effects quickly become very important, and flow separation can occur. The Northrop Grumman USAF B-2 Bomber encountered a nonlinear aeroelastic Residual Pitch Oscillation (RPO) under these conditions. Simulation studies were performed with the Computational Aeroelasticity Program-Transonic Small Disturbance, Viscous (CAPTSDv) computer program to evaluate its ability to predict these nonlinear aeroelastic responses. Open and closed loop simulations were performed to assess the participation of the flight control system. Control, system actuator hysteresis characteristics were modeled and found to be a significant participant in the RPO phenomenon. Simulations were also performed for varying Mach numbers and altitudes to establish the stability boundaries and compare with flight test data. These CAPTSDv simulations compared well with flight data and revealed many potential further modeling enhancements.
Wolf, S. W. D.
Work has continued with the low speed self streamlining wind tunnel (SSWT) using the NACA 0012-64 airfoil in an effort to explain the discrepancies between the NASA Langley low turbulence pressure tunnel (LTPT) and SSWT results obtained with the airfoil stalled. Conventional wind tunnel corrections were applied to straight wall SSWT airfoil data, to illustrate the inadequacy of standard correction techniques in circumstances of high blockage. Also one SSWT test was re-run at different air speeds to investigate the effects of such changes on airfoil data and wall contours. Mechanical design analyses for the transonic self streamlining wind tunnel (TSWT) were completed by the application of theoretical airfoil flow field data to the elastic beam and streamline analysis. The control system for the transonic facility is outlined.
Kato, Shoji; Honma, Fumio; Matsumoto, Ryoji
Viscous instability of the transonic region of the conventional geometrically thin alpha-type accretion disks is examined analytically. For simplicity, isothermal disks and isothermal perturbations are assumed. It is found that when the value of alpha is larger than a critical value the disk is unstable against two types of perturbations. One is local propagating perturbations of inertial acoustic waves. Results suggest the possibility that unstable perturbations develop to overstable global oscillations which are restricted only in the innermost region of the disk. The other is standing growing perturbations localized just at the transonic point. The cause of these instabilities is that the azimuthal component of the Lagrangian velocity variation associated with the perturbations becomes in phase with the variation of the viscous stress force. Because of this phase matching work is done on perturbations, and they are amplified.
Cook, W. J.; Chaney, M. J.; Presley, L. L.; Chapman, G. T.
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.
Tomek, W. G.; Hall, R. M.; Wahls, R. A.; Luckring, J. M.; Owens, L. R.
A wind tunnel test of a generic fighter configuration was tested in the National Transonic Facility through a cooperative agreement between NASA Langley Research Center and McDonnell Douglas. The primary purpose of the test was to assess Reynolds number scale effects on a thin-wing, fighter-type configuration up to full-scale flight conditions (that is, Reynolds numbers of the order of 60 million). The test included longitudinal and lateral/directional studies at subsonic and transonic conditions across a range of Reynolds numbers from that available in conventional wind tunnels to flight conditions. Results are presented for three Mach numbers (0.6, 0.8, and 0.9) and three configurations: (1) Fuselage/Wing; (2) Fuselage/Wing/Centerline Vertical Tail/Horizontal Tail; and (3) Fuselage/Wing/Trailing-Edge Extension/Twin Vertical Tails. Reynolds number effects on the longitudinal aerodynamic characteristics are presented herein.
Chattot, J. J.
A finite difference code for predicting the high speed flow over the advancing helicopter rotor is presented. The code solves the low frequency, transonic small disturbance equation and is suitable for modeling the effects of advancing blade unsteadiness on blades of nearly arbitrary planform. The method employs a quasi-conservative mixed differencing scheme and solves the resulting difference equations by an alternating direction scheme. Computed results showed good agreement with experimental blade pressure data and illustrate some of the effects of varying the rotor planform. The flow unsteadiness is shown to be an indispensible part of a transonic solution. Close to the tip at high advance ratio, cross flow effects can significantly affect the solution.
Vadyak, J.; Atta, E. H.
A highly efficient computer analysis was developed for predicting transonic nacelle/inlet flowfields. This algorithm can compute the three dimensional transonic flowfield about axisymmetric (or asymmetric) nacelle/inlet configurations at zero or nonzero incidence. The flowfield is determined by solving the full-potential equation in conservative form on a body-fitted curvilinear computational mesh. The difference equations are solved using the AF2 approximate factorization scheme. This report presents a discussion of the computational methods used to both generate the body-fitted curvilinear mesh and to obtain the inviscid flow solution. Computed results and correlations with existing methods and experiment are presented. Also presented are discussions on the organization of the grid generation (NGRIDA) computer program and the flow solution (NACELLE) computer program, descriptions of the respective subroutines, definitions of the required input parameters for both algorithms, a brief discussion on interpretation of the output, and sample cases to illustrate application of the analysis.
Kemp, W. B., Jr.
A method for assessing the wall interference in transonic two dimensional wind tunnel test was developed and implemented in a computer program. The method involves three successive solutions of the transonic small disturbance potential equation to define the wind tunnel flow, the perturbation attriburable to the model, and the equivalent free air flow around the model. Input includes pressure distributions on the model and along the top and bottom tunnel walls which are used as boundary conditions for the wind tunnel flow. The wall induced perturbation fields is determined as the difference between the perturbation in the tunnel flow solution and the perturbation attributable to the model. The methodology used in the program is described and detailed descriptions of the computer program input and output are presented. Input and output for a sample case are given.
Igoe, William B.
Dynamic measurements of fluctuating static pressure levels were taken with flush-mounted, high-frequency response pressure transducers at 11 locations in the circuit of the National Transonic Facility (NTF) across the complete operating range of this wind tunnel. Measurements were taken at test-section Mach numbers from 0.1 to 1.2, at pressures from 1 to 8.6 atm, and at temperatures from ambient to -250 F, which resulted in dynamic flow disturbance measurements at the highest Reynolds numbers available in a transonic ground test facility. Tests were also made by independent variation of the Mach number, the Reynolds number, or the fan drive power while the other two parameters were held constant, which for the first time resulted in a distinct separation of the effects of these three important parameters.
Wright, L. C.; Vitale, N. G.; Ware, T. C.; Erwin, J. R.
A high-tip-speed, low-loading transonic fan stage was designed to deliver an overall pressure ratio of 1.5 with an adiabatic efficiency of 86 percent. The design flow per unit annulus area is 42.0 pounds per square foot. The fan features a hub/tip ratio of 0.46, a tip diameter of 28.74 in. and operates at a design tip speed of 1600 fps. For these design conditions, the rotor blade tip region operates with supersonic inlet and supersonic discharge relative velocities. A sophisticated quasi-three-dimensional characteristic section design procedure was used for the all-supersonic sections and the inlet of the midspan transonic sections. For regions where the relative outlet velocities are supersonic, the blade operates with weak oblique shocks only.
Schewe, G.; Mai, H.; Dietz, G.
This paper presents flutter and forced oscillation experiments in a transonic wind tunnel. For an aeroelastic supercritical 2-D airfoil configuration we studied typical transonic phenomena in as pure a form as possible. Various manifestations of small-amplitude limit cycle oscillations were observed for different flow conditions as well as coexisting limit cycles. We demonstrated how very small control forces were sufficient to excite or suppress flutter oscillations. Limit cycle oscillations occurred under free and forced turbulent boundary layer transition in a perforated wall test-section. Flutter calculations based on experimental aerodynamic forces yield stability limits which show good agreement with directly measured experimental flutter values. The results indicate that flow separation at the trailing edge, and the interactions between the shock and the marginal region of separated flow beneath it, may be responsible for limiting the amplitude of the observed limit cycle oscillations.
Reed, T. D.; Moretti, P. M.
The laminar boundary layer on a 10 degree cone in a transonic wind tunnel was studied. The inviscid flow and boundary layer development were simulated by computer programs. The effects of pitch and yaw angles on the boundary layer were examined. Preston-tube data, taken on the boundary-layer-transition cone in the NASA Ames 11 ft transonic wind tunnel, were used to develope a correlation which relates the measurements to theoretical values of laminar skin friction. The recommended correlation is based on a compressible form of the classical law-of-the-wall. The computer codes successfully simulates the laminar boundary layer for near-zero pitch and yaw angles. However, in cases of significant pitch and/or yaw angles, the flow is three dimensional and the boundary layer computer code used here cannot provide a satisfactory model. The skin-friction correlation is thought to be valid for body geometries other than cones.
Bennett, Robert M.; Dansberry, Bryan E.; Farmer, Moses G.; Eckstrom, Clinton V.; Seidel, David A.; Rivera, Jose A., Jr.
Transonic shock boundary layer oscillations occur on rigid models over a small range of Mach numbers on thick circular-arc airfoils. Extensive tests and analyses of this phenomena have been made in the past but essentially all of them were for rigid models. A simple flexible wing model with an 18 pct. circular arc airfoil was constructed and tested in the Langley Transonic Dynamics Tunnel to study the dynamic characteristics that a wing might have under these circumstances. In the region of shock boundary layer oscillations, buffeting of the first bending mode was obtained. This mode was well separated in frequency from the shock boundary layer oscillations. A limit cycle oscillation was also measured in a third bending like mode, involving wind vertical bending and splitter plate motion, which was in the frequency range of the shock boundary layer oscillations. Several model configurations were tested, and a few potential fixes were investigated.
Hooker, John R.; Burner, Alpheus W.; Valla, Robert
A computational method for accurately predicting the static aeroelastic deformations of typical transonic transport wind tunnel models is described. The method utilizes a finite element method (FEM) for predicting the deformations. Extensive calibration/validation of this method was carried out using a novel wind-off wind tunnel model static loading experiment and wind-on optical wing twist measurements obtained during a recent wind tunnel test in the National Transonic Facility (NTF) at NASA LaRC. Further validations were carried out using a Navier-Stokes computational fluid dynamics (CFD) flow solver to calculate wing pressure distributions about several aeroelastically deformed wings and comparing these predictions with NTF experimental data. Results from this aeroelastic deformation method are in good overall agreement with experimentally measured values. Including the predicted deformations significantly improves the correlation between CFD predicted and experimentally measured wing & pressures.
Vincenti, Walter G; Wagoner, Cleo B
A theoretical study is described of the aerodynamic characteristics at small angle of attack of a thin, double-wedge profile in the range of supersonic flight speed in which the bow wave is detached. The analysis is carried out within the framework of the transonic (nonlinear) small-disturbance theory, and the effects of angle of attack are regarded as a small perturbation on the flow previously calculated at zero angle. The mixed flow about the front half of the profile is calculated by relaxation solution of a suitably defined boundary-value problem for transonic small-disturbance equation in the hodograph plane (i.e., the Tricomi equation). The purely supersonic flow about the rear half is found by an extension of the usual numerical method of characteristics. Analytical results are also obtained, within the framework of the same theory, for the range of speed in which the bow wave is attached and the flow is completely supersonic.
Sydnor, George H.; Bhatia, Ram; Krattiger, Hansueli; Mylius, Justus; Schafer, D.
In September 1995, a project was initiated to replace the existing drive line at NASA's most unique transonic wind tunnel, the National Transonic Facility (NTF), with a single 101 MW synchronous motor driven by a Load Commutated Inverter (LCI). This Adjustable Speed Drive (ASD) system also included a custom four-winding transformer, harmonic filter, exciter, switch gear, control system, and feeder cable. The complete system requirements and design details have previously been presented and published , as well as the commissioning and acceptance test results . The NTF was returned to service in December 1997 with the new drive system powering the fan. Today, this installation still represents the world s largest horizontal single motor/drive combination. This paper describes some significant events that occurred with the drive system during the first 15 years of service. These noteworthy issues are analyzed and root causes presented. Improvements that have substantially increased the long term viability of the system are given.
Gloss, B. B.; Nystrom, D.
The National Transonic Facility (NTF), a fan-driven, transonic, pressurized, cryogenic wind tunnel, will operate over the Mach number range of 0.10 to 1.20 with stagnation pressures varying from 1.00 to about 8.8 atm and stagnation temperatures varying from 77 to 340 K. The NTF is cooled to cryogenic temperatures by the injection of liquid nitrogen into the tunnel stream with gaseous nitrogen as the test gas. The NTF can also operate at ambient temperatures using a conventional chilled water heat exchanger with air on nitrogen as the test gas. The methods used in estimating the fan pressure ratio requirements are described. The estimated NTF operating envelopes at Mach numbers from 0.10 to 1.20 are presented.
Keener, Earl R.
Tabulations and plots are presented of boundary-layer velocity and flow-direction surveys from wind-tunnel tests of a large-scale (0.90 m semi-span) model of the NASA/Lockheed Wing C. This wing is a generic, transonic, supercritical, highly three-dimensional, low-aspect-ratio configuration designed with the use of a three-dimensional, transonic full-potential-flow wing code (FLO22). Tests were conducted at the design angle of attack of 5 deg over a Mach number range from 0.25 to 0.96 and a Reynolds number range of 3.4x10 to the 6th power. Wing pressures were measured at five span stations, and boundary-layer surveys were measured at the midspan station. The data are presented without analysis.
Whitlow, W., Jr.; Seidel, D. A.
Nonreflecting far-field boundary conditions that are consistent with the complete transonic small-disturbance (TSD) equations are derived. They are implemented in a new code for solving the complete TSD equation and are tested for a harmonically oscillating NACA 64A010 airfoil in transonic flow and for a flat-plate airfoil with a pulse in the angle of attack. Using the new boundary conditions on a relatively small grid, solutions for the airfoil that are obtained that agree with large-grid calculations, resulting in a 44 percent savings in computer time. Frequency responses for the flat plate show that most of the disturbances incident on the computational boundaries are absorbed by the boundary conditions.
Yu, Y. H.; Kittleson, J. K.; Becker, F.
Holographic interferometry and computer-assisted tomography (CAT) are used to determine the transonic velocity field of a model rotor blade in hover. A pulsed ruby laser recorded 40 interferograms with a 2-ft-diam view field near the model rotor-blade tip operating at a tip Mach number of 0.90. After digitizing the interferograms and extracting fringe-order functions, the data are transferred to a CAT code. The CAT code then calculates the perturbation velocity in seeral planes above the blade surface. The values from the holography-CAT method compare favorably with previously obtained numerical computations in most locations near the blade tip. The results demonstrate the technique's potential for three-dimensional transonic rotor flow studies.
Kittleson, John K.; Yu, Yung H.
Holographic interferometry and computerized aided tomography (CAT) are used to determine the transonic velocity field of a model rotor blade in hover. A pulsed ruby laser recorded 40 interferograms with a 2 ft dia view field near the model rotor blade tip operating at a tip Mach number of 0.90. After digitizing the interferograms and extracting the fringe order functions, the data are transferred to a CAT code. The CAT code then calculates the perturbation velocity in several planes above the blade surface. The values from the holography-CAT method compare favorably with previously obtained numerical computations in most locations near the blade tip. The results demonstrate the technique's potential for three dimensional transonic rotor flow studies.
Kittleson, J. K.; Yu, Y. H.; Becker, F.
Holographic interferometry and computer-assisted tomography (CAT) are used to determine the transonic flow field of a model rotor blade in hover. A pulsed ruby laser records 40 interferograms with a 61 cm-diam view field near the model rotor-blade tip operating at a tip Mach number of 0.90. After digitizing the interferograms and extracting fringe-order functions, the data are transferred to a CAT code. The CAT code then calculates pressure coefficients in several planes above the blade surface. The values from the holography-CAT method compare favorably with previously obtained numerical computations and laser velocimeter measurements at most locations near the blade tip. The results demonstrate the technique's potential for three-dimensional transonic rotor flow studies.
Yu, Y. H.; Kittleson, J. K.; Becker, F.
Holographic interferometry and computer-assisted tomography (CAT) are used to determine the transonic velocity field of a model rotor blade in hover. A pulsed ruby laser recorded 40 interferograms with a 2-ft-diam view field near the model rotor-blade tip operating at a tip Mach number of 0.90. After digitizing the interferograms and extracting fringe-order functions, the data are transferred to a CAT code. The CAT code then calculates the perturbation velocity in several planes above the blade surface. The value from the holography-CAT method compare favorably with previously obtained numerical computations in most locations near the blade tip. The results demonstrate the technique's potential for three-dimensional transonic rotor flow studies.
Tuzla, K.; Russell, D. A.; Wai, J. C.
Nonlifting 10% biconvex airfoils are mounted in a 30 x 40 cm Ludwieg-tube-driven transonic test-section and the flow field recorded with a holographic interferometer. Nitrogen, argon, and carbon dioxide are used as the principal test gases. Experiments are conducted with Reynolds number based on chord of (0.5-3.5) x 10 to the 6th with Mach numbers of 0.70, 0.75, and 0.80. Supporting calculations use inviscid transonic small-disturbance and full-potential computer codes coupled with simple integral boundary-layer modeling. Systematic studies show that significant gamma-effects can occur due to shock-induced separation.
Chambers, Joseph R.; Hall, Robert M.
This paper presents the results of a survey of past experiences with uncommanded lateral-directional motions at transonic speeds during specific military aircraft programs. The effort was undertaken to provide qualitative and quantitative information on past airplane programs that might be of use to the participants in the joint NASA/Navy/Air Force Abrupt Wing Stall (AWS) Program. The AWS Program was initiated because of the experiences of the F/A-l8E/F development program, during which unexpected, severe wing-drop motions were encountered by preproduction aircraft at transonic conditions. These motions were judged to be significantly degrading to the primary mission requirements of the aircraft. Although the problem was subsequently solved for the production version of the F/A-l8E/F, a high-level review panel emphasized the poor understanding of such phenomena and issued a strong recommendation to: "Initiate a national research effort to thoroughly and systematically study the wing drop phenomena." A comprehensive, cooperative NASA/Navy/Air Force AWS Program was designed to respond to provide the required technology requirements. As part of the AWS Program, a work element was directed at a historical review of wing-drop experiences in past aircraft development programs at high subsonic and transonic speeds. In particular, information was requested regarding: specific aircraft configurations that exhibited uncommanded motions and the nature of the motions; geometric characteristics of the air- planes; flight conditions involved in occurrences; relevant data, including wind-tunnel, computational, and flight sources; figures of merit used for analyses; and approaches used to alleviate the problem. An attempt was also made to summarize some of the more important lessons learned from past experiences, and to recommend specific research efforts. In addition to providing technical information to assist the AWS research objectives, the study produced fundamental
Chambers, Joseph R.; Hall, Robert M.
This paper presents the results of a survey of past experiences with uncommanded lateral-directional motions at transonic speeds during specific military aircraft programs. The effort was undertaken to provide qualitative and quantitative information on past airplane programs that might be of use to the participants in the joint NASA/Navy/Air Force Abrupt Wing Stall (AWS) Program. The AWS Program was initiated because of the experiences of the F/A-18E/F development program, during which unexpected, severe wing-drop motions were encountered by preproduction aircraft at transonic conditions. These motions were judged to be significantly degrading to the primary mission requirements of the aircraft. Although the problem was subsequently solved for the production version of the F/A-l8E/F, a high-level review panel emphasized the poor understanding of such phenomena and issued a strong recommendation to: Initiate a national research effort to thoroughly and systematically study the wing drop phenomena. A comprehensive, cooperative NASA/Navy/Air Force AWS Program was designed to respond to provide the required technology requirements. As part of the AWS Program, a work element was directed at a historical review of wing-drop experiences in past aircraft development programs at high subsonic and transonic speeds. In particular, information was requested regarding: specific aircraft configurations that exhibited uncommanded motions and the nature of the motions; geometric characteristics of the air- planes; flight conditions involved in occurrences; relevant data, including wind-tunnel, computational, and flight sources; figures of merit used for analyses; and approaches used to alleviate the problem. An attempt was also made to summarize some of the more important lessons learned from past experiences, and to recommend specific research efforts. In addition to providing technical information to assist the AWS research objectives, the study produced fundamental information
Bobbitt, Percy J.; Harvey, William D.; Harris, Charles D.; Brooks, Cuyler W., Jr.
An account is given of the considerations involved in selecting the NASA-Langley transonic pressure tunnel's design and test parameters, as well as its liner and a swept wing for laminar flow control (LFC) experimentation. Attention is given to the types and locations of the instrumentation employed. Both slotted and perforated upper surfaces were tested with partial- and full-chord suction; representative results are presented for all.
Introduction In transonic turbomachinery airfoils, both shock waves and wakes appear at the trailing edge, giving rise stator-rotor interactions that...Journal of Turbomachinery , Vol. 129, pp. 740- 749.  Ifeachor, E. C. and Jervis, B. W., 1996, Digital Signal Processing, Addison- Wesley, New York...Composite Fan Stator,” ASME Journal of Turbomachinery , Vol. 127, pp. 412-421 (Also ASME Paper No. GT2004-53492).  Doorly, D. J. and Oldfield, M. L
Kilgore, W. Allen; Balakrishna, S.
The 0.3 m Transonic Cryogenic Tunnel (TCT) microcomputer based controller has been operating for several thousand hours in a safe and efficient manner. A complete listing is provided of the source codes for the tunnel controller and tunnel simulator. Included also is a listing of all the variables used in these programs. Several changes made to the controller are described. These changes are to improve the controller ease of use and safety.
Mercer, C. E.; Burley, J. R., II
The Langley 16-Foot Transonic Tunnel was used to evaluate the performance characteristics of a coannular plug nozzle at static conditions (Mach number of 0) and at Mach numbers from 0.65 to 1.20. Jet total pressure ratio was varied from 1.0 (jet off) to 10.0. Thirty-seven configurations generated by the combination of three geometric variables - plug angle, shroud boattail length (fixed exit radius), and shroud extension length - were tested.
Newman, P. A.; Davis, R. M.
The input parameters are presented for a computer program which performs calculations for inviscid isentropic transonic flow over three dimensional airfoils with straight leading edges. The free stream Mach number is restricted only by the isentropic assumption. Weak shock waves are automatically located where they occur in the flow. The finite difference form of the full equation for the velocity potential is solved by the method of relaxation, after the flow exterior to the airfoil is mapped to the upper half plane.
Mercer, Charles E.; Berrier, Bobby L.; Capone, Francis J.; Grayston, Alan M.
The equations used by the 16-Foot Transonic Wind Tunnel in the data reduction programs are presented in nine modules. Each module consists of equations necessary to achieve a specific purpose. These modules are categorized in the following groups: (1) tunnel parameters; (2) jet exhaust measurements; (3) skin friction drag; (4) balance loads and model attitudes calculations; (5) internal drag (or exit-flow distribution); (6) pressure coefficients and integrated forces; (7) thrust removal options; (8) turboprop options; and (9) inlet distortion.
Paryz, Roman W.
Several upgrade projects have been completed or are just getting started at the NASA Langley Research Center National Transonic Facility. These projects include a new high capacity semi-span balance, model dynamics damping system, semi-span model check load stand, data acquisition system upgrade, facility automation system upgrade and a facility reliability assessment. This presentation will give a brief synopsis of each of these efforts.
Takashima, K.; Sawada, H.; Aoki, T.
The facilities for aerodynamic testing of airplane models at transonic speeds and high Reynolds numbers are surveyed. The need for high Reynolds number testing is reviewed, using some experimental results. Some approaches to high Reynolds number testing such as the cryogenic wind tunnel, the induction driven wind tunnel, the Ludwieg tube, the Evans clean tunnel and the hydraulic driven wind tunnel are described. The level of development of high Reynolds number testing facilities in Japan is discussed.
Springer, A. M.
A proposed wing-body reusable launch vehicle was tested in the NASA Marshall Space Flight Center's 14 x 14-inch trisonic wind tunnel during the winter of 1994. This test resulted in the vehicle's subsonic and transonic, Mach 0.3 to 1.96, longitudinal and lateral aerodynamic characteristics. The effects of control surface deflections on the basic vehicle's aerodynamics, including a body flap, elevons, ailerons, and tip fins, are presented.
Springer, Anthony M.
A proposed wing body reusable launch vehicle was tested in the Marshall Space Flight Center (MSFC) 14 x 14 inch trisonic wind tunnel during the winter of 1994. this test resulted in the vehicle's subsonic and transonic (Mach 03 to 1.96) longitudinal and lateral aerodynamic characteristics. The effects of control surface deflections on the basic vehicle aerodynamics including a body flap, elevons, ailerons, and tip fins are presented.
Bartels, Robert E.
Three-dimensional transonic flow over a delta wing is investigated using several turbulence models. The performance of linear eddy viscosity models and an explicit algebraic stress model is assessed at the start of vortex flow, and the results compared with experimental data. To assess the effect of transition location, computations that either fix transition aft of the leading edge or are fully turbulent are performed. These computations show that grid resolution, transition location and turbulence model significantly affect the 3D flowfield.
Bobbitt, P. J.; Carter, J. E.
Requirements for flow quality in the National Transonic Facility are explored. Viscous flow effects of concern to theoreticians are discussed. Experiments outlined for theory validation in the facility include validating high aspect ratio wing-body combination; low aspect ratio moderately swept wing; low aspect ratio highly swept wing; high lift systems on high aspect ration wings; Reynolds number scaling; dynamic shock- boundary layer interaction; and the effect of R and M on dynamic stall.
Campbell, Richard L.; Smith, Leigh A.
A method has been developed for designing airfoils and wings at transonic speeds. It utilizes a hybrid design algorithm in an iterative predictor/corrector approach, alternating between analysis code and a design module. This method has been successfully applied to a variety of airfoil and wing design problems, including both transport and highly-swept fighter wing configurations. An efficient approach to viscous airfoild design and the effect of including static aeroelastic deflections in the wing design process are also illustrated.
Carlson, Leland A.
Research conducted during the period from July 1991 through December 1992 is covered. A method based upon the quasi-analytical approach was developed for computing the aerodynamic sensitivity coefficients of three dimensional wings in transonic and subsonic flow. In addition, the method computes for comparison purposes the aerodynamic sensitivity coefficients using the finite difference approach. The accuracy and validity of the methods are currently under investigation.
ION BETWEEN THE ROTOR f AtZ STATOR OF A TRANSONIC XCMPRESSCR F. NEIl-OFF EXOTEH, INC. 3935 SEAMN AVE, SUITE 0 COTRACTR 1985OR FINAL RERT FOR PERID MAY...406.24. UNCLASSIFIED TABLE OF CONTETS Page CF’r~ FIGURES . . . . .. ..... i L ST OF TN =E . ........... 0.......... v I. Introduction...At 0% of design speed . ........ ...................... 7 iv %* LISr CF TABLES Page I. 1)Rta from file 795608 . . 38 II. Data from file 7977 ....5 7
Suggested fluid mechanics research to be conducted in the National Transonic Facility include: wind tunnel calibration; flat plate skin friction, flow visualization and measurement techniques; leading edge separation; high angle of attack separation; shock-boundary layer interaction; submarine shapes; low speed studies of cylinder normal to flow; and wall interference effects. These theoretical aerodynamic investigations will provide empirical inputs or validation data for computational aerodynamics, and increase the usefulness of existing wind tunnels.
Silva, Walter A.
The presentation begins with a brief description of the motivation and approach that has been taken for this research. This will be followed by a description of the Volterra Theory of Nonlinear Systems and the CAP-TSD code which is an aeroelastic, transonic CFD (Computational Fluid Dynamics) code. The application of the Volterra theory to a CFD model and, more specifically, to a CAP-TSD model of a rectangular wing with a NACA 0012 airfoil section will be presented.
engine, so the steam -induced stall characteristic of its compressor must be well understood to help prevent any catastrophic failures of the aircraft...INVESTIGATION OF HIGH-PRESSURE STEAM INDUCED STALL OF A TRANSONIC ROTOR by Joesph J. Koessler June 2007 Thesis Advisor: Garth V. Hobson...2007 3. REPORT TYPE AND DATES COVERED Master’s Thesis 4. TITLE AND SUBTITLE Experimental Investigation of High-Pressure Steam Induced Stall of a
Hah, Chunill; Voges, Melanie; Mueller, Martin; Schiffer, Heinz-Peter
In the present study, unsteady flow behavior in a modern transonic axial compressor rotor is studied in detail with large eddy simulation (LES) and particle image velocimetry (PIV). The main purpose of the study is to advance the current understanding of the flow field near the blade tip in an axial transonic compressor rotor near the stall and peak-efficiency conditions. Flow interaction between the tip leakage vortex and the passage shock is inherently unsteady in a transonic compressor. Casing-mounted unsteady pressure transducers have been widely applied to investigate steady and unsteady flow behavior near the casing. Although many aspects of flow have been revealed, flow structures below the casing cannot be studied with casing-mounted pressure transducers. In the present study, unsteady velocity fields are measured with a PIV system and the measured unsteady flow fields are compared with LES simulations. The currently applied PIV measurements indicate that the flow near the tip region is not steady even at the design condition. This self-induced unsteadiness increases significantly as the compressor rotor operates near the stall condition. Measured data from PIV show that the tip clearance vortex oscillates substantially near stall. The calculated unsteady characteristics of the flow from LES agree well with the PIV measurements. Calculated unsteady flow fields show that the formation of the tip clearance vortex is intermittent and the concept of vortex breakdown from steady flow analysis does not seem to apply in the current flow field. Fluid with low momentum near the pressure side of the blade close to the leading edge periodically spills over into the adjacent blade passage. The present study indicates that stall inception is heavily dependent on unsteady behavior of the flow field near the leading edge of the blade tip section for the present transonic compressor rotor.
Variable Definition ..... .............. ... 16 d. Transonic Performance Optimization .. ......... ... 17 e. Wind Tunnel Test...44 36 Wing Section Shape Model for Wing Design Optimization . . .. 45 37 Wind Tunnel Model in AEDC PWT 16T Tunnel ............. .... 46 38 Test...FPE-designed wing was selected as the final C-1418/AC2 wing design. 4. WIND TUNNEL TEST a. Test Facility The design verification wind tunnel tests were
The thin-layer Navier-Stokes equations are coupled with a zonal scheme (or domain-decomposition method) to develop the Transonic Navier-Stokes (TNS) wing-alone code. The TNS has a total of 4 zones and is extended to a total of 16 zones for the wing-fuselage version of the code. Results are compared on the Cray X-MP-48 and compared with experimental data.
Young, C. P., Jr.; Gerringer, A. H.; Brooks, T. G.; Berry, R. F., Jr.
The feasibility of making weld repairs on heavy section 9% nickel steel forgings such as those being manufactured for the National Transonic Facility fan disk and fan drive shaft components was evaluated. Results indicate that 9% nickel steel in heavy forgings has very good weldability characteristics for the particular weld rod and weld procedures used. A comparison of data for known similar work is included.
Grenon, R.; Desopper, A.; Sides, J.
The experimental results of steady and unsteady pressure measurements, carried out in subsonic and transonic flow on a 16 percent relative thickness supercritical aerofoil, equipped with a trailing edge flap involving 25 percent of the chord, in a sinusoidal motion are given. These experimental results are compared with those obtained by various methods of steady and unsteady inviscid flow calculations. Some calculation results in which viscous effects have been taken into account, for both steady and unsteady flows, are also presented.
This report represents a general theory applicable to axial, radial, and mixed flow turbomachines operating at subsonic and supersonic speeds with a finite number of blades of finite thickness. References reflect the evolution of computational methods used, from the inception of the theory in the 50's to the high-speed computer era of the 90's. Two kinds of relative stream surfaces, S(sub 1) and S(sub 2), are introduced for the purpose of obtaining a three-dimensional flow solution through the combination of two-dimensional flow solutions. Nonorthogonal curvilinear coordinates are used for the governing equations. Methods of computing transonic flow along S(sub 1) and S(sub 2) stream surfaces are given for special cases as well as for fully three-dimensional transonic flows. Procedures pertaining to the direct solutions and inverse solutions are presented. Information on shock wave locations and shapes needed for computations are discussed. Experimental data from a Deutsche Forschungs- und Versuchsanstalt fur Luft- und Raumfahrt e.V. (DFVLR) rotor and from a Chinese Academy of Sciences (CAS) transonic compressor rotor are compared with the computed flow properties.
Owens, D. Bruce; Capone, Francis J.; Brandon, Jay M.; Cunningham, Kevin; Chambers, Joseph R.
Transonic free-to-roll and static wind tunnel tests for four military aircraft - the AV-8B, the F/A-18C, the preproduction F/A-18E, and the F-16C - have been analyzed. These tests were conducted in the NASA Langley 16-Foot Transonic Tunnel as a part of the NASA/Navy/Air Force Abrupt Wing Stall Program. The objectives were to evaluate the utility of the free-to-roll test technique as a tool for predicting areas of significant uncommanded lateral motions and for gaining insight into the wing-drop and wing-rock behavior of military aircraft at transonic conditions. The analysis indicated that the free-to-roll results had good agreement with flight data on all four models. A wide range of motions - limit cycle wing rock, occasional and frequent damped wing drop/rock and wing rock divergence - were observed. The analysis shows the effects that the static and dynamic lateral stability can have on the wing drop/rock behavior. In addition, a free-to-roll figure of merit was developed to assist in the interpretation of results and assessment of the severity of the motions.
Wu, J. M.; Collins, F. G.
The free-stream interference caused by the flow through the slotted walls of the test sections of transonic wind tunnels has continuously a problem in transonic tunnel testing. The adaptive-wall transonic tunnel is designed to actively control the near-wall boundary conditions by sucking or blowing through the wall. In order to make the adaptive-wall concept work, parameters for computational boundary conditions must be known. These parameters must be measured with sufficient accuracy to allow numerical convergence of the flow field computations and must be measured in an inviscid region away from the model that is placed inside the wind tunnel. The near-wall flow field was mapped in detail using a five-port cone probe that was traversed in a plane transverse to the free-stream flow. The initial experiments were made using a single slot and recent measurements used multiple slots, all with the tunnel empty. The projection of the flow field velocity vectors on the transverse plane revealed the presence of a vortex-like flow with vorticity in the free stream. The current research involves the measurement of the flow field above a multislotted system with segmented plenums behind it, in which the flow is controlled through several plenums simultaneously. This system would be used to control a three-dimensional flow field.
The aerodynamic measurements in conventional wind tunnels usually suffer from the interference effects of the sting supporting the model and the test section walls. These effects are particularly severe in the transonic regime. Sting interference effects can be overcome through the Magnetic Suspension technique. Wall effects can be alleviated by: testing airfoils in conventional, ventilated tunnels at relatively small model to tunnel size ratios; treatment of the tunnel wall boundary layers; or by utilization of the Adaptive Wall Test Section (AWTS) concept. The operating capabilities and results from two of the foremost two-dimensional, transonic, AWTS facilities in existence are assessed. These facilities are the NASA 0.3-Meter Transonic Cryogenic Tunnel and the ONERA T-2 facility located in Toulouse, France. In addition, the results derived from the well known conventional facility, the NAE 5 ft x 5 ft Canadian wind tunnel will be assessed. CAST10/D0A2 Airfoil results will be used in all of the evaluations.
VanZante, Dale E.; Strazisar, Anthony J.; Wood, Jerry R,; Hathaway, Michael D.; Okiishi, Theodore H.
The tip clearance flows of transonic compressor rotors are important because they have a significant impact on rotor and stage performance. While numerical simulations of these flows are quite sophisticated. they are seldom verified through rigorous comparisons of numerical and measured data because these kinds of measurements are rare in the detail necessary to be useful in high-speed machines. In this paper we compare measured tip clearance flow details (e.g. trajectory and radial extent) with corresponding data obtained from a numerical simulation. Recommendations for achieving accurate numerical simulation of tip clearance flows are presented based on this comparison. Laser Doppler Velocimeter (LDV) measurements acquired in a transonic compressor rotor, NASA Rotor 35, are used. The tip clearance flow field of this transonic rotor was simulated using a Navier-Stokes turbomachinery solver that incorporates an advanced k-epsilon turbulence model derived for flows that are not in local equilibrium. Comparison between measured and simulated results indicates that simulation accuracy is primarily dependent upon the ability of the numerical code to resolve important details of a wall-bounded shear layer formed by the relative motion between the over-tip leakage flow and the shroud wall. A simple method is presented for determining the strength of this shear layer.
Ivanco, Thomas G.
NASA Langley Research Center's Transonic Dynamics Tunnel (TDT) is the world's most capable aeroelastic test facility. Its large size, transonic speed range, variable pressure capability, and use of either air or R-134a heavy gas as a test medium enable unparalleled manipulation of flow-dependent scaling quantities. Matching these scaling quantities enables dynamic similitude of a full-scale vehicle with a sub-scale model, a requirement for proper characterization of any dynamic phenomenon, and many static elastic phenomena. Select scaling parameters are presented in order to quantify the scaling advantages of TDT and the consequence of testing in other facilities. In addition to dynamic testing, the TDT is uniquely well-suited for high risk testing or for those tests that require unusual model mount or support systems. Examples of recently conducted dynamic tests requiring unusual model support are presented. In addition to its unique dynamic test capabilities, the TDT is also evaluated in its capability to conduct aerodynamic performance tests as a result of its flow quality. Results of flow quality studies and a comparison to a many other transonic facilities are presented. Finally, the ability of the TDT to support future NASA research thrusts and likely vehicle designs is discussed.
Viscous flow past a wind tunnel model of a 65-degree swept angle Delta wing at transonic speeds is being studied. The model was tested in the 8-foot cryogenic transonic wind tunnel at the National Transonic Facility. Aerodynamic forces and wing surface pressure data were obtained at various angles of attack, Mach numbers, and Reynold's numbers for four different leading edges of the wing. The objectives of the present investigation are: (1) to perform numerical modeling of the flow around the wing; (2) to validate the experimental data with a Navier-Stokes computational fluid dynamics code and vice versa; (3) to investigate the effects of the sting mount of the wing; (4) to evaluate the effects of leading edge radius on the flow; and (5) to explain the Reynold's number effect as indicated by the test data. Several computer programs were developed to define the surfaces of the wing, the four leading edges, and the sting mount. Based on these geometric databases, the surface grids of a single-block computational domain was generated interactively on the IRIS workstation using the GRIDGEN2D module of GRIDGEN. To refine the grids and to avoid excessive loss of grid points due to collapsed edges, a 9-block computational domain containing approximately 750,000 grid points was developed with the GRIDBLOCK module to replace the single-block grid.
Kuhlman, John M.; Brown, Christopher K.
Computational designs were performed for three different low aspect ratio wing planforms fitted with nonplanar winglets; one of the three configurations was selected to be constructed as a wind tunnel model for testing in the NASA LaRC 8-foot transonic pressure tunnel. A design point of M = 0.8, C(sub L) is approximate or = to 0.3 was selected, for wings of aspect ratio equal to 2.2, and leading edge sweep angles of 45 deg and 50 deg. Winglet length is 15 percent of the wing semispan, with a cant angle of 15 deg, and a leading edge sweep of 50 deg. Winglet total area equals 2.25 percent of the wing reference area. The design process and the predicted transonic performance are summarized for each configuration. In addition, a companion low-speed design study was conducted, using one of the transonic design wing-winglet planforms but with different camber and thickness distributions. A low-speed wind tunnel model was constructed to match this low-speed design geometry, and force coefficient data were obtained for the model at speeds of 100 to 150 ft/sec. Measured drag coefficient reductions were of the same order of magnitude as those predicted by numerical subsonic performance predictions.
Carter, Melissa B.; Vicroy, Dan D.; Patel, Dharmendra
The Blended-Wing-Body (BWB) concept has shown substantial performance benefits over conventional aircraft configuration with part of the benefit being derived from the absence of a conventional empennage arrangement. The configuration instead relies upon a bank of trailing edge devices to provide control authority and augment stability. To determine the aerodynamic characteristics of the aircraft, several wind tunnel tests were conducted with a 2% model of Boeing's BWB-450-1L configuration. The tests were conducted in the NASA Langley Research Center's National Transonic Facility and the Arnold Engineering Development Center s 16-Foot Transonic Tunnel. Characteristics of the configuration and the effectiveness of the elevons, drag rudders and winglet rudders were measured at various angles of attack, yaw angles, and Mach numbers (subsonic to transonic speeds). The data from these tests will be used to develop a high fidelity simulation model for flight dynamics analysis and also serve as a reference for CFD comparisons. This paper provides an overview of the wind tunnel tests and examines the effects of Reynolds number, Mach number, pitch-pause versus continuous sweep data acquisition and compares the data from the two wind tunnels.
Fukushima, Yuma; Kawai, Soshi
In this study, we conduct the wall-modeled large-eddy simulation (LES) of transonic buffet phenomena over the OAT15A supercritical airfoil at high Reynolds number. The transonic airfoil buffet involves shock-turbulent boundary layer interactions and shock vibration associated with the flow separation downstream of the shock wave. The wall-modeled LES developed by Kawai and Larsson PoF (2012) is tuned on the K supercomputer for high-fidelity simulation. We first show the capability of the present wall-modeled LES on the transonic airfoil buffet phenomena and then investigate the detailed flow physics of unsteadiness of shock waves and separated boundary layer interaction phenomena. We also focus on the sustaining mechanism of the buffet phenomena, including the source of the pressure waves propagated from the trailing edge and the interactions between the shock wave and the generated sound waves. This work was supported in part by MEXT as a social and scientific priority issue to be tackled by using post-K computer. Computer resources of the K computer was provided by the RIKEN Advanced Institute for Computational Science (Project ID: hp150254).
Goodliff, Scott L.; Balakrishna, Sundareswara; Butler, David; Cagle, C. Mark; Chan, David; Jones, Gregory S.; Milholen, William E., II
The National Transonic Facility is a transonic pressurized cryogenic facility. The development of the high Reynolds number semi-span capability has advanced over the years to include transonic active flow control and powered testing using the sidewall model support system. While this system can be used in total temperatures down to -250Â F for conventional unpowered configurations, it is limited to temperatures above -60Â F when used with powered models that require the use of the high-pressure air delivery system. Thermal instabilities and non-repeatable mechanical arrangements revealed several data quality shortfalls by the force and moment measurement system. Recent modifications to the balance cavity recirculation system have improved the temperature stability of the balance and metric model-to-balance hardware. Changes to the mechanical assembly of the high-pressure air delivery system, particularly hardware that interfaces directly with the model and balance, have improved the repeatability of the force and moment measurement system. Drag comparisons with the high-pressure air system removed will also be presented in this paper.
Lepicovsky, Jan; Chima, Rodrick V.; Jett, Thomas A.; Bencic, Timothy J.; Weiland, Kenneth E.
An extensive study into the nature of the separated flows on the suction side of modem transonic fan airfoils at high incidence is described in the paper. Suction surface.flow separation is an important flow characteristic that may significantly contribute to stall flutter in transonic fans. Flutter in axial turbomachines is a highly undesirable and dangerous self-excited mode of blade oscillations that can result in high cycle fatigue blade failure. The study basically focused on two visualization techniques: surface flow visualization using dye oils, and schlieren (and shadowgraph) flow visualization. The following key observations were made during the study. For subsonic inlet flow, the flow on the suction side of the blade is separated over a large portion of the blade, and the separated area increases with increasing inlet Mach number. For the supersonic inlet flow condition, the flow is attached from the leading edge up to the point where a bow shock from the upper neighboring blade hits the blade surface. Low cascade solidity, for the subsonic inlet flow, results in an increased area of separated flow. For supersonic flow conditions, a low solidity results in an improvement in flow over the suction surface. Finally, computational results modeling the transonic cascade flowfield illustrate our ability to simulate these flows numerically.
Milholen, William E., II; Owens, Lewis R.
The effect of discrete contour bumps on reducing the transonic drag at off-design conditions on an airfoil have been examined. The research focused on fully-turbulent flow conditions, at a realistic flight chord Reynolds number of 30 million. State-of-the-art computational fluid dynamics methods were used to design a new baseline airfoil, and a family of fixed contour bumps. The new configurations were experimentally evaluated in the 0.3-m Transonic Cryogenic Tunnel at the NASA Langley Research center, which utilizes an adaptive wall test section to minimize wall interference. The computational study showed that transonic drag reduction, on the order of 12% - 15%, was possible using a surface contour bump to spread a normal shock wave. The computational study also indicated that the divergence drag Mach number was increased for the contour bump applications. Preliminary analysis of the experimental data showed a similar contour bump effect, but this data needed to be further analyzed for residual wall interference corrections.
Hanson, P. W.
The characteristics and capabilities of the two tunnels, that relate to studies in the fields of aeroelasticity and unsteady aerodynamics are discussed. Scaling considerations for aeroelasticity and unsteady aerodynamics testing in the two facilities are reviewed, and some of the special features (or lack thereof) of the Langley Research Center Transonic Dynamics Tunnel (TDT) and the National Transonic Facility (NTF) that will weigh heavily in any decisions conducting a given study in the two tunnels are discussed. For illustrative purposes a fighter and a transport airplane are scaled for tests in the NTF and in the TDT, and the resulting model characteristics are compared. The NTF was designed specifically to meet the need for higher Reynolds number capability for flow simulation in aerodynamic performance testing of aircraft designs. However, the NTF can be a valuable tool for evaluating the severity of Reynolds number effects in the areas of dynamic aeroelasticity and unsteady aerodynamics. On the other hand, the TDT was constructed specifically for studies and tests in the field of aeroelasticity. Except for tests requiring the Reynolds number capability of NTF, the TDT will remain the primary facility for tests of dynamic aeroelasticity and unsteady aerodynamics.
Carlson, J. R.; Compton, W. B., III
A wind tunnel investigation was conducted to determine the aerodynamic interference associated with the installation of a long duct, flow-through nacelle on a straight unswept untapered supercritical wing. Experimental data was obtained for the verification of computational prediction techniques. The model was tested in the 16-Foot Transonic Tunnel at Mach numbers from 0.20 to 0.875 and at angles of attack from about 0 deg to 5 deg. The results of the investigation show that strong viscous and compressibility effects are present at the transonic Mach numbers. Numerical comparisons show that linear theory is adequate for subsonic Mach number flow prediction, but is inadequate for prediction of the extreme flow conditions that exist at the transonic Mach numbers.
Grossman, B.; Moretti, G.
A computer program to predict the inviscid, transonic flow field about isolated nacelles was developed. The problem was to be formulated to solve Euler's equations without any approximation (such as small disturbances) and hence the terminology exact solution. The flow field was complicated by the presence of imbedded shock waves, an engine-inlet interface, and exhaust plumes. Furthermore, the transonic nacelles of interest had a very slender but blunt cowl lip. This created two distinct length scales, the length of the nacelle and the cowl lip radius that can differ by several orders of magnitude. These aspects of the flow field presented many numerical difficulties. The approach to the problem was to calculate the nacelle flow field using the method of time-dependent computations (TDC). Although at the time of the issuance of this contract, other approaches to transonic flow calculations existed, it was felt that TDC offered the most effective means of meeting the goals of the contract.
Warner, R. W.; Wilcox, P. R.
Free-oscillation damping measurements at hypersonic and transonic Mach numbers are presented for the lower pitch flap of the M2-F2 reentry vehicle. For the flow and model conditions tested, the damping measurements indicate the absence of hypersonic buzz instability for flap rotation frequencies of 47.3 Hz, 153 Hz, and 360 Hz, the presence of transonic buzz for 47.3 Hz flap, and the absence of transonic buzz for a 115 Hz flap. There are not enough flap damping data for an error estimate based on repeatability, but a partial damping calibration is presented in which an analog computer simulation of random flap response for a known flap damping is fed into the autocorrelation computer and filter combination used for most of the damping measurements.
Ruhlin, C. L.; Rauch, F. J., Jr.; Waters, C.
The model was a 1/6.5-size, semipan version of a wing proposed for an executive-jet-transport airplane. The model was tested with a normal wingtip, a wingtip with winglet, and a normal wingtip ballasted to simulate the winglet mass properties. Flutter and aerodynamic data were acquired at Mach numbers (M) from 0.6 to 0.95. The measured transonic flutter speed boundary for each wingtip configuration had roughly the same shape with a minimum flutter speed near M=0.82. The winglet addition and wingtip mass ballast decreased the wing flutter speed by about 7 and 5 percent, respectively; thus, the winglet effect on flutter was more a mass effect than an aerodynamic effect.
Ladson, Charles L.; Hill, Acquilla S.; Johnson, William G., Jr.
Tests were conducted in the 2-D test section of the Langley 0.3-meter Transonic Cryogenic Tunnel on a NACA 0012 airfoil to obtain aerodynamic data as a part of the Advanced Technology Airfoil Test (ATAT) program. The test program covered a Mach number range of 0.30 to 0.82 and a Reynolds number range of 3.0 to 45.0 x 10 to the 6th power. The stagnation pressure was varied between 1.2 and 6.0 atmospheres and the stagnation temperature was varied between 300 K and 90 K to obtain these test conditions. Tabulated pressure distributions and integrated force and moment coefficients are presented as well as plots of the surface pressure distributions. The data are presented uncorrected for wall interference effects and without analysis.
Seidel, B. S.; Matwey, M. D.; Adamczyk, J. J.
In the present paper, a semi-actuator-disk theory is reviewed that was developed previously for the distorted inflow to a single-stage axial-flow compressor. Flow distortion occurs far upstream; it may be a distortion in stagnation temperature, stagnation pressure, or both. Losses, quasi-steady deviation angles, and reference incidence correlations are included in the analysis, and both subsonic and transonic relative Mach numbers are considered. The theory is compared with measurements made in a transonic fan stage, and a parameter study is carried out to determine the influence of solidity on the attenuation of distortions in stagnation pressure and stagnation temperature.
Hallissy, J. B.; Ayers, T. G.
An investigation was conducted in the NASA Langley 8-foot transonic pressure tunnel at Mach numbers from 0.60 to 0.975 with a variable-wing-sweep airplane model in order to evaluate a series of wings designed to demonstrate the maneuver potential of the supercritical airfoil concept. Both conventional and supercritical wing designs for several planform configurations were investigated with wing sweep angles from 16.0 deg to 72.5 deg, depending on Mach number and wing configuration. The supercritical wing configuration showed significant improvements over the conventional configurations in drag-divergence Mach number and in drag level at transonic maneuver conditions.
Batina, John T.
A time-accurate approximate-factorization (AF) algorithm is described for solution of the three-dimensional unsteady transonic small-disturbance equation. The AF algorithm consists of a time-linearization procedure coupled with a subiteration technique. The algorithm is the basis for the Computational Aeroelasticity Program-Transonic Small Disturbance (CAP-TSD) computer code, which was developed for the analysis of unsteady aerodynamics and aeroelasticity of realistic aircraft configurations. The paper describes details on the governing flow equations and boundary conditions, with an emphasis on documenting the finite-difference formulas of the AF algorithm.
Bonhaus, Daryl L.; Wornom, Stephen F.
Two codes which solve the 3-D Thin Layer Navier-Stokes (TLNS) equations are used to compute the steady state flow for two test cases representing typical finite wings at transonic conditions. Several grids of C-O topology and varying point densities are used to determine the effects of grid refinement. After a description of each code and test case, standards for determining code efficiency and accuracy are defined and applied to determine the relative performance of the two codes in predicting turbulent transonic wing flows. Comparisons of computed surface pressure distributions with experimental data are made.
Dippold, Vance F., III; Zaman, Khairul B. M. Q.
Hot-wire and acoustic measurements were taken for a round convergent nozzle and a round convergent-divergent (C-D) nozzle at a jet Mach number of 0.61. The C-D nozzle had a design Mach number of 2.2. Compared to the convergent nozzle jet flow, the Mach 2.2 nozzle jet flow produced excess broadband noise (EBBN). It also produced a transonic resonance tone at 1200 Herz. Computational simulations were performed for both nozzle flows. A steady Reynolds-Averaged Navier-Stokes simulation was performed for the convergent nozzle jet flow. For the Mach 2.2 nozzle flow, a steady RANS simulation, an unsteady RANS (URANS) simulation, and an unsteady Detached Eddy Simulation (DES) were performed. The RANS simulation of the convergent nozzle showed good agreement with the hot-wire velocity and turbulence measurements, though the decay of the potential core was over-predicted. The RANS simulation of the Mach 2.2 nozzle showed poor agreement with the experimental data, and more closely resembled an ideally-expanded jet. The URANS simulation also showed qualitative agreement with the hot-wire data, but predicted a transonic resonance at 1145 Herz. The DES showed good agreement with the hot-wire velocity and turbulence data. The DES also produced a transonic tone at 1135 Herz. The DES solution showed that the destabilization of the shock-induced separation region inside the nozzle produced increased levels of turbulence intensity. This is likely the source of the EBBN.
Abu-Mostafa, A. S.; Reed, T. D.
Preston-tube measurements obtained on the Arnold Engineering Development Center (AEDC) Transition Cone have been correlated with theoretical skin friction coefficients in transitional and turbulent flow. This has been done for the NASA Ames 11-Ft Transonic Wind Tunnel (11 TWT) and flight tests. The developed semi-empirical correlations of Preston-tube data have been used to derive a calibration procedure for the 11 TWT flow quality. This procedure has been applied to the corrected laminar data, and an effective freestream unit Reynolds number is defined by requiring a matching of the average Preston-tube pressure in flight and in the tunnel. This study finds that the operating Reynolds number is below the effective value required for a match in laminar Preston-tube data. The distribution of this effective Reynolds number with Mach number correlates well with the freestream noise level in this tunnel. Analyses of transitional and turbulent data, however, did not result in effective Reynolds numbers that can be correlated with background noise. This is a result of the fact that vorticity fluctuations present in transitional and turbulent boundary layers dominate Preston-tube pressure fluctuations and, therefore, mask the tunnel noise eff ects. So, in order to calibrate the effects of noise on transonic wind tunnel tests only laminar data should be used, preferably at flow conditions similar to those in flight tests. To calibrate the effects of transonic wind-tunnel noise on drag measurements, however, the Preston-tube data must be supplemented with direct measurements of skin friction.
Hah, Chunill; Rabe, Douglas; Scribben, Angie
In the study presented, unsteady flow interaction between an ultra-compact inlet and a transonic fan stage is investigated. Future combat aircraft engines require ultra-compact inlet ducts as part of an integrated, advanced propulsion system to improve air vehicle capability and effectiveness to meet future mission needs. The main purpose of the current study is to advance the understanding of the flow interaction between a modern ultra-compact inlet and a transonic fan for future design applications. Many experimental/ analytical studies have been reported on the aerodynamics of compact inlets in aircraft engines. On the other hand, very few studies have been reported on the effects of flow distortion from these inlets on the performance of the following fan/compressor stages. The primary goal of the study presented is to investigate how flow interaction between an ultra-compact inlet and a transonic compressor influence the operating margin of the compressor. Both Unsteady Reynolds-averaged Navier-Stokes (URANS) and Large Eddy Simulation (LES) approaches are used to calculate the unsteady flow field, and the numerical results are used to study the flow interaction. The present study indicates that stall inception of the following compressor stage is affected directly based on how the distortion pattern evolves before it interacts with the fan/compressor face. For the present compressor, the stall initiates at the tip section with clean inlet flow and distortion pattern away from the casing itself seems to have limited impacts on the stall inception of the compressor. A counter-rotating swirl, which is generated due to flow separation inside the s-shaped compact duct, generates an increased flow angle near the blade tip. This increased flow angle near the rotor tip due to the secondary flow from the counter-rotating vortices is the primary reason for the reduced compressor stall margin.
Ueno, Atsushi; Suzuki, Kojiro
For the success of hypersonic vehicles, their shape must be optimized to achieve a high lift-to-drag ratio as well as a low aerodynamic heating rate in the hypersonic regime. In addition, the transonic lift-to-drag ratio must also be optimized to realize quick acceleration to the hypersonic cruise speed. The three-dimensional lift-to-drag ratio can be improved even by the two-dimensional section shape (i.e., airfoil) optimization in the region where the sweep back angle is small. Here, prior to three-dimensional shape optimization, a study is done to optimize airfoils of hypersonic vehicles based on these three parameters. At optimization, the hypersonic lift-to-drag ratio is maximized while the transonic lift-to-drag ratio and the aerodynamic heating rate are constrained. The optimum lift coefficient for hypersonic cruise at the maximum lift-to-drag ratio is investigated. The relation between the leading edge radius, which determines the aerodynamic heating rate, and the hypersonic lift-to-drag ratio is also investigated. Results show that to improve the hypersonic lift-to-drag ratio, the airfoil thickness around the leading edge should be small as long as an appropriate compromise with the transonic lift-to-drag ratio is achieved. Results also show that the optimum lift coefficient for hypersonic cruise is much lower than that for typical supersonic vehicles. Small cruise lift coefficient suggests that the wing loading of a hypersonic vehicle should be small. The leading edge radius should be determined by a compromise between the hypersonic lift-to-drag ratio and leading edge heating. Airfoil optimization can provide an appropriate initial guess of the three-dimensional optimum shape. By using an appropriate initial guess, the computation time of the three-dimensional shape optimization is expected to be reduced.
We have developed a method to model the effects of an engine nacelle on the transonic flow about a swept wing. The nacelle is modelled by adding an inhomogeneous term, which represents a source distribution, to the potential equation, which is solved by relaxation. Our method is an extension of the design code FL22INV of Bauer, Garabedian, and McFadden, which was based on Caughy and Jameson's FLO22 and on the oblique wing code FLO16. In order to obtain a sharply defined obstacle using a source distribution on a finite difference grid, we must use a small grid size near the source. We accomplish this through a mapping of the flow domain to the computational domain. We thus obtain a concentrated source distribution which, for transonic Mach numbers, will cause high speeds and cavitation. To overcome this problem we use a constant coefficient equation for the velocity potential near the source. Once the flow has been computed, the size and position of the nacelle are found by tracing streamlines. We have used our code to study the effect of the presence of an engine nacelle on the shock waves which form above a swept wing in transonic flight. It was found that sizeable shocks appear on wings that were previously nearly shockless. By applying the free boundary redesign technique of FL22INV we were able to minimize the wave drag due to these shocks. The results of these studies are included along with a listing of our new FORTRAN code NACROSS, in which an effort has been made to improve the implementation. 26 refs., 14 figs.
An extensive experimental study into the nature of the separated flows on the blade suction surface of modern transonic fans is described in this paper. The study was a subtask of a larger experimental effort focused on blade flutter excited by flow separation in the blade tip region. The tip sections of airfoils on transonic fan blades are designed for precompression and consequently they differ from sections on the rest of the blade. The blade tip section was modeled by a low aspect ratio blade and therefore most of the blade tested was exposed to the secondary flow effects. The aim of this work was to supply reliable data on flow separation on transonic fan blades for validation of future analytical studies. The experimental study focused on two visualization techniques: surface flow visualization using dye oils and schlieren (and shadowgraph) flow visualization. The following key observations were made during the study. For subsonic inlet flow, the flow on the suction surface of the blade was separated over a large portion of the blade, and the separated area increased with increasing inlet Mach number. For the supersonic inlet flow condition, the flow was attached from the leading edge up to the point where a bow shock from the upper neighboring blade imposed on the blade surface. Downstream, there was a separated flow region in which air flowed in the direction opposite the inlet flow. Finally, past the separated flow region, the flow reattached to the blade surface. For subsonic inlet flow, the low cascade solidity resulted in an increased area of separated flow. For supersonic flow conditions, the low solidity resulted in an improvement in flow over the suction surface.
Snow, W. L.; Burner, A. W.; Goad, W. K.
The optical quality of gas in a cryogenic wind tunnel was determined by observing Sayce targets through different pathlengths of the medium. The data were used to determine the square wave response of the test gas. At conditions corresponding to 15 times ambient density, considerable decrease in response to higher spatial frequencies was noted even in the absence of flow. Under flow conditions, vibrations further degraded the response. The results are interpreted in terms of possible photogrammetric approaches to measure model deformation in large cryogenic facilities such as the National Transonic Facility.
This paper reports the result of an analysis of wind tunnel data acquired in support of the Facility Analysis Verification & Operational Reliability (FAVOR) project. The analysis uses methods referred to collectively at Langley Research Center as the Modern Design of Experiments (MDOE). These methods quantify the total variance in a sample of wind tunnel data and partition it into explained and unexplained components. The unexplained component is further partitioned in random and systematic components. This analysis was performed on data acquired in similar wind tunnel tests executed in four different U.S. transonic facilities. The measurement environment of each facility was quantified and compared.
A study has been made of possible ways to improve the performance of the Langley Research Center's Transonic Dynamics Tunnel (TDT). The major effort was directed toward obtaining increased dynamic pressure in the Mach number range from 0.8 to 1.2, but methods to increase Mach number capability were also considered. Methods studied for increasing dynamic pressure capability were higher total pressure, auxiliary suction, reducing circuit losses, reduced test medium temperature, smaller test section and higher molecular weight test medium. Increased Mach number methods investigated were nozzle block inserts, variable geometry nozzle, changes in test section wall configuration, and auxiliary suction.
Solutions of steady transonic flow past a two-dimensional airfoil are obtained from a singular integro-differential equation which involves a tangential derivative of the perturbation velocity potential. Subcritical flows are solved by taking central differences everywhere. For supercritical flows with shocks, central differences are taken in subsonic flow regions and backward differences in supersonic flow regions. The method is applied to a nonlifting parabolic-arc airfoil and to a lifting NACA 0012 airfoil. Results compare favorably with those of finite-difference schemes.
Urasek, D. C.; Janetzke, D. C.
The design and experimental performance of a 20-inch diameter tandem-bladed axial-flow transonic compressor rotor is presented. Radial surveys were made of the flow conditions. At design speed the peak efficiency was 0.88 and occurred at an equivalent weight flow of 63 pounds per second. At peak efficiency the total pressure and total temperature ratios were 1.77 and 1.20, respectively. The stall margin at design speed was 10 percent based on weight flows and total pressure ratios at peak efficiency and near stall.
Carlson, Leland A.
The three dimensional quasi-analytical sensitivity analysis and the ancillary driver programs are developed needed to carry out the studies and perform comparisons. The code is essentially contained in one unified package which includes the following: (1) a three dimensional transonic wing analysis program (ZEBRA); (2) a quasi-analytical portion which determines the matrix elements in the quasi-analytical equations; (3) a method for computing the sensitivity coefficients from the resulting quasi-analytical equations; (4) a package to determine for comparison purposes sensitivity coefficients via the finite difference approach; and (5) a graphics package.
Dowell, E. H.
A simplified theory of the dynamic motion of aerofoils of finite thickness in transonic flow is presented which excludes the effect of shock waves on the aerofoil itself and, thus, is restricted to free stream Mach numbers equal to unity or above. Numerical examples are analyzed for two-dimensional steady and unsteady (including transient) aerofoil motion and three-dimensional steady and unsteady flow over delta wings. The effects of flow separation and improvements in Bernoulli's equation and the surface boundary condition are briefly discussed.
Boyden, R. P.; Brooks, C. W., Jr.; Davenport, E. E.
Static and dynamic stability tests were made of a finned projectile configuration with the aft-mounted fins arranged in a cruciform pattern. The tests were made at free stream Mach numbers of 0.7, 0.9, 1.1, and 1.2 in the Langley 8-foot transonic pressure tunnel. Some of the parameters measured during the tests were lift, drag, pitching moment, pitch damping, and roll damping. Configurations tested included the body with undeflected fins, the body with various fin deflections for control, and the body with fins removed. Theoretical estimates of the stability derivatives were made for the fins on configuration.
Melson, N. D.; Keller, J. D.
In this paper, the authors report on some of their experiences modifying two three-dimensional transonic flow programs (FLO22 and FLO27) for use on the NASA Langley Research Center CYBER 203. Both of the programs discussed were originally written for use on serial machines. Several methods were attempted to optimize the execution of the two programs on the vector machine, including: (1) leaving the program in a scalar form (i.e., serial computation) with compiler software used to optimize and vectorize the program, (2) vectorizing parts of the existing algorithm in the program, and (3) incorporating a new vectorizable algorithm (ZEBRA I or ZEBRA II) in the program.
Miser, James W; Stewart, Warner L; Monroe, Daniel E
The subject turbine was investigated to determine the effect of high rotor pressure-surface diffusion on turbine performance. A comparison of the subject turbine with the most efficient transonic turbine in the present series of investigations showed that the efficiency of the subject turbine was almost as high, the suction-surface diffusion parameter was about the same, and the solidity was reduced by 36 percent. Because the loss per blade increased greatly with an increase in pressure-surface diffusion, the latter is also considered to be an important design consideration.
Roberts, W. B.; Crouse, J. E.; Sandercock, D. M.
Experimental data from 10 transonic fan rotors were used to correlate losses created by part-span dampers located near the midchord position on the rotor blades. The design tip speed of these rotors varied from 419 to 425 m/sec, and the design pressure ratio varied from 1.6 to 2.0. Additional loss caused by the dampers for operating conditions between 50 and 100 percent of design speed were correlated with relevant aerodynamic and geometric parameters. The resulting correlation predicts the variation of total-pressure-loss coefficient in the damper region to a good approximation.
27709 tINDER CONTRACT DAAG29-78-C-0004 PREPARED BY: 0 ANALYTICAL METHODS , INC. 2047 - 152ND AVENUE N.E. REDMOND., WASHINGTON 98052 ( (206) 643-9090...Analytical Methods , Inc. AREA & WORK UNIT NUMBERS 2047 - 152nd Avenue N.E. Redmond, Washington 98052 11. CONTROLIN(’ OFFICE NAME AND ADDRESS 12. REPORT...analysis 2& ATaRAcT (entiare - ,efeea ad t neweesen and idatly by block inbet) A calculation method for transonic separated flow about two-dimensional air
Oberkampf, W. L.
This report describes a theoretical method for the prediction of fin forces and moments on bodies at high angle of attack in subsonic and transonic flow. The body is assumed to be a circular cylinder with cruciform fins (or wings) of arbitrary planform. The body can have an arbitrary roll (or bank) angle, and each fin can have individual control deflection. The method combines a body vortex flow model and lifting surface theory to predict the normal force distribution over each fin surface. Extensive comparisons are made between theory and experiment for various planform fins. A description of the use of the computer program that implements the method is given.
Keller, J. D.; Jameson, A.
An explicit method for solving the transonic small-disturbance potential equation is presented. This algorithm, which is suitable for the new vector-processor computers such as the CDC STAR-100, is compared to successive line over-relaxation (SLOR) on a simple test problem. The convergence rate of the explicit scheme is slower than that of SLOR, however, the efficiency of the explicit scheme on the STAR-100 computer is sufficient to overcome the slower convergence rate and allow an overall speedup compared to SLOR on the CYBER 175 computer.
Kittleson, J. K.
A technique that uses holographic interferometry to record the first interferograms of the flow near a hovering transonic rotor blade is presented. A pulsed ruby laser is used to record interferograms of a 2 ft diam field of view near a rotor tip operating at a tip Mach number of 0.90. Several interferograms, recorded along planes perpendicular to the rotor's tip path plane at various azimuth angles around the flow, are presented. These interferograms yield quantitative information about shock structure and location, flow separation, and radiated noise that will help helicopter researchers understand the complexities of the flow around high speed rotor blades and thus improve performance and reduce noise.
Mann, M. J.; Langhans, R. A.
The effects of wing trailing-edge control surfaces on the static transonic aerodynamic characteristics of a transport configuration with a supercritical wing were studied. The configuration was tested with both an area-ruled fuselage and a cylindrical fuselage. The Mach number range was from 0.80 to 0.96 and the angle of attack range was from -1 deg to 12 deg. The Reynolds number was 1,580,000 based on the mean aerodynamic chord. Tabular data are presented.
Henderson, William P.; Burley, James R., II
An investigation has been conducted in the Langley 16-Foot Transonic Tunnel to determine the effects on empennage arrangement on single-engine nozzle/afterbody static pressures. Tests were done at Mach numbers from 0.60 to 1.20, nozzle pressure ratios from 1.0 (jet off) to 8.0. and angles of attack from -3 to 9 deg (at jet off conditions), depending on Mach number. Three empennage arrangements (aft, staggered, and forward) were investigated. Extensive measurements were made of static pressure on the nozzle/afterbody in the vicinity of the tail surfaces.
The linear transonic perturbation integral equation previously derived for nonlifting airfoils is formulated for lifting cases. In order to treat shock wave motions, a strained coordinate system is used in which the shock location is invariant. The tangency boundary conditions are either formulated using the thin airfoil approximation or by using the analytic continuation concept. A direct numerical solution to this equation is derived in contrast to the iterative scheme initially used, and results of both lifting and nonlifting examples indicate that the method is satisfactory.
Bland, Samuel R.
Calculation of unsteady flow phenomena requires careful attention to the numerical treatment to the governing partial differential equations. The personal computer provides a convenient and useful tool for the development of meshes, algorithms, and boundary conditions needed to provide time accurate solution of these equations. The one-dimensional equation considered provides a suitable model for the study of wave propagation in the equations of transonic small disturbance potential flow. Numerical results for effects of mesh size, extent, and stretching, time step size, and choice of far-field boundary conditions are presented. Analysis of the discretized model problem supports these numerical results. Guidelines for suitable mesh and time step choices are given.
Thibodeaux, J. J.
Guidelines and suggestions substantiated by real-time simulation data to ensure optimum time and energy use of injected liquid nitrogen for cooling the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT) are presented. It is directed toward enabling operators and researchers to become cognizant of criteria for using the 0.3-m TCT in an energy- or time-efficient manner. Operational recommendations were developed based on information collected from a validated simulator of the 0.3-m TCT and experimental data from the tunnel. Results and trends, however, can be extrapolated to other similarly constructed cryogenic wind tunnels.
Ruo, S. Y.
A computer program has been developed to account approximately for the effects of finite wing thickness in the transonic potential flow over an oscillating wing of finite span. The program is based on the original sonic-box program of Rodemich and Andrew, and accounts for the nonuniform flow caused by finite thickness by application of the local linearization concept. A brief description of each subroutine is given, and the method of input is shown in detail. A sample problem as well as a complete listing of the computer program are presented.
Wright, Ray H; Ritchie, Virgil S; Pearson, Albin O
A large wind tunnel, approximately 8 feet in diameter, has been converted to transonic operation by means of slots in the boundary extending in the direction of flow. The usefulness of such a slotted wind tunnel, already known with respect to the reduction of the subsonic blockage interference and the production of continuously variable supersonic flows, has been augmented by devising a slot shape with which a supersonic test region with excellent flow quality could be produced. Experimental locations of detached shock waves ahead of axially symmetric bodies at low supersonic speeds in the slotted test section agreed satisfactorily with predictions obtained by use of existing approximate methods.
Roberts, W. B.
The design-point losses caused by part-span dampers (PSD) were correlated for 21 transonic axial flow fan rotors that had tip speeds varying from 350 to 488 meters per second and design pressure ratios of 1.5 to 2.0. For these rotors a correlation using mean inlet Mach number at the damper location, along with relevant geometric and aerodynamic loading parameters, predicts the variation of total pressure loss coefficient in the region of the damper to a good approximation.
Spreiter, John R; Alksne, Alberta Y
A method is presented for the approximate solution of the nonlinear equations of transonic flow theory. Solutions are found for two-dimensional flows at a Mach number of 1 and for purely subsonic and purely supersonic flows. Results are obtained in closed analytic form for a large and significant class of nonlifting airfoils. At a Mach number of 1 general expressions are given for the pressure distribution on an airfoil of specified geometry and for the shape of an airfoil having a prescribed pressure distribution. Extensive comparisons are made with available data, particularly for a Mach number of 1, and with existing solutions.
Volpe, G.; Grossman, B.
The multiphase effort in the development of tools for the analysis and design of two-element airfoil systems, that is, airfoils with a slat or a flap at transonic speeds is described. The first phase involved the development of a method to compute the inviscid flow over such configurations. In the second phase the inviscid code was coupled to a boundary layer calculation program in order to compute the loss in performance due to viscous effects. An inverse code that constructs the airfoil system corresponding to a desired pressure distribution is described.
Spreiter, John R; Alksne, Alberta Y
A method is presented for the approximate solution of the nonlinear equations transonic flow theory. Solutions are found for two-dimensional flows at a Mach number of 1 and for purely subsonic and purely supersonic flows. Results are obtained in closed analytic form for a large and significant class of nonlifting airfoils. At a Mach number of 1 general expressions are given for the pressure distribution on an airfoil of specified geometry and for the shape of an airfoil having a prescribed pressure distribution. Extensive comparisons are made with available data, particularly for a Mach number of 1, and with existing solutions.
Lawing, P. L.; Kilgore, R. A.
The model building, development, and testing experience gained during 8 years of operation of the 0.3-m Transonic Cryogenic Tunnel (TCT) is summarized. The summary is divided into four portions: (1) models tested in the 0.3-m TCT's original octagonal test section; (2) models tested in the present two dimensional test section; (3) models tested as a part of tunnel calibration and the development of advanced technology airfoils; and (4) development of a new way to construct two dimensional airfoil models. Design requirements imposed on the models by high Reynolds number testing at cryogenic temperatures are reviewed.
Farrell, C.; Adamczyk, J.
The three-dimensional flow in a turbomachinery blade row was approximated by correcting for streamtube convergence and radius change in the throughflow direction. The method is a fully conservative solution of the full potential equation incorporating the finite volume technique on body fitted periodic mesh, with an artificial density imposed in the transonic region to insure stability and the capture of shock waves. Comparison of results for several supercritical blades shows good agreement with their hodograph solutions. Other calculations for these profiles as well as standard NACA blade sections indicate that this is a useful scheme analyzing both the design and off-design performance of turbomachinery blading.
Carlson, Leland A.
A direct-inverse technique and computer program called TAMSEP that can be used for the analysis of the flow about airfoils at subsonic and low transonic freestream velocities is presented. The method is based upon a direct-inverse nonconservative full potential inviscid method, a Thwaites laminar boundary layer technique, and the Barnwell turbulent momentum integral scheme; and it is formulated using Cartesian coordinates. Since the method utilizes inverse boundary conditions in regions of separated flow, it is suitable for predicting the flow field about airfoils having trailing edge separated flow under high lift conditions. Comparisons with experimental data indicate that the method should be a useful tool for applied aerodynamic analyses.
Investigation of downstream boundary effects on the frequency of self-excited oscillations in two-dimensional, separated transonic diffuser flows were conducted numerically by solving the compressible, Reynolds-averaged, thin-layer Navier-Stokes equation with two equation turbulence models. It was found that the flow fields are very sensitive to the location of the downstream boundary. Extension of the diffuser downstream boundary significantly reduces the frequency and amplitude of oscillations for pressure, velocity, and shock. The existence of a suction slot in the experimental setpup obscures the physical downstream boundary and therefore presents a difficulty for quantitative comparisons between computation and experiment.
Hsieh, T.; Coakley, T. J.
A numerical investigation of two-dimensional unsteady boundary layer in a transonic diffuser flow with self-excited oscillations and strong flow separation by solving the compressible, Reynolds-averaged, thin-layer Navier-Stokes equations with two-equations turbulence model is described. Three different meshes with constant streamwise mesh distribution and varying vertical mesh distribution were used. Results obtained indicate that a refinement of mesh studied here has minimal effect on the mean boundary layer flow but significantly increases the amplitude of oscillation of all flow variables. Comparisons of unsteady wall pressure, velocity profile, terminal shock, and separation pocket among computations and with experiment are presented.
Lewis, M. C.
The original flexible top and bottom walls of the Transonic Self-Streamlining Wind Tunnel (TSWT), at the University of Southampton, have been replaced with new walls featuring a larger number of static pressure tappings and detailed mechanical improvements. This report describes the streamling method, results, and conclusions of a series of tests aimed at defining sets of aerodynamically straight wall contours for the new flexible walls. This procedure is a necessary prelude to model testing. The quality of data obtained compares favorably with the aerodynamically straight data obtained with the old walls. No operational difficulties were experienced with the new walls.
Budovsky, A. D.; Sidorenko, A. A.; Polivanov, P. A.; Vishnyakov, O. I.; Maslov, A. A.
The study deals with experimental investigation of unsteady features of separated flow on a profiled bump in transonic environment. The experiments were conducted in T-325 wind tunnel of ITAM for the following flow conditions: P0 = 1 bar, T0 = 291 K. The base flow around the model was studied by schlieren visualization, steady and unsteady wall pressure measurements and PIV. The experimentally data obtained using PIV are analyzed by Proper Orthogonal Decomposition (POD) technique to investigate the underlying unsteady flow organization, as revealed by the POD eigenmodes. The data obtained show that flow pulsations revealed upstream and downstream of shock wave are correlated and interconnected.
Bland, Samuel R.
Calculation of unsteady flow phenomena requires careful attention to the numerical treatment of the governing partial differential equations. The personal computer provides a convenient and useful tool for the development of meshes, algorithms, and boundary conditions needed to provide time accurate solution of these equations. The one-dimensional equation considered provides a suitable model for the study of wave propagation in the equations of transonic small disturbance potential flow. Numerical results for effects of mesh size, extent, and stretching, time step size, and choice of far-field boundary conditions are presented. Analysis of the discretized model problem supports these numerical results. Guidelines for suitable mesh and time step choices are given.
Melson, N. D.
The analysis and incorporation into a multigrid scheme of several vectorizable algorithms are discussed. Von Neumann analyses of vertical line, horizontal line, and alternating direction ZEBRA algorithms were performed; and the results were used to predict their multigrid damping rates. The algorithms were then successfully implemented in a transonic conservative full-potential computer program. The convergence acceleration effect of multiple grids is shown and the convergence rates of the vectorizable algorithms are compared to the convergence rates of standard successive line overrelaxation (SLOR) algorithms.
Florance, James R.; Rivera, Jose A., Jr.
The Transonic Dynamics Tunnel(TDT) was recalibrated due to the conversion of the heavy gas test medium from R-12 to R-134a. The objectives of the tests were to determine the relationship between the free-stream Mach number and the measured test section Mach number, and to quantify any necessary corrections. Other tests included the measurement of pressure distributions along the test-section walls, test-section centerline, at certain tunnel stations via a rake apparatus, and in the tunnel settling chamber. Wall boundary layer, turbulence, and flow angularity measurements were also performed. This paper discusses the determination of sidewall Mach number distributions.
Carlson, K. D.
A direct-inverse technique and computer program called TAMSEP that can be sued for the analysis of the flow about airfoils at subsonic and low transonic freestream velocities is presented. The method is based upon a direct-inverse nonconservative full potential inviscid method, a Thwaites laminar boundary layer technique, and the Barnwell turbulent momentum integral scheme; and it is formulated using Cartesian coordinates. Since the method utilizes inverse boundary conditions in regions of separated flow, it is suitable for predicing the flowfield about airfoils having trailing edge separated flow under high lift conditions. Comparisons with experimental data indicate that the method should be a useful tool for applied aerodynamic analyses.
Thurman, Douglas; Flegel, Ashlie; Giel, Paul
Constant temperature hotwire anemometry data were acquired to determine the inlet turbulence conditions of a transonic turbine blade linear cascade. Flow conditions and angles were investigated that corresponded to the take-off and cruise conditions of the Variable Speed Power Turbine (VSPT) project and to an Energy Efficient Engine (EEE) scaled rotor blade tip section. Mean and turbulent flowfield measurements including intensity, length scale, turbulence decay, and power spectra were determined for high and low turbulence intensity flows at various Reynolds numbers and spanwise locations. The experimental data will be useful for establishing the inlet boundary conditions needed to validate turbulence models in CFD codes.
Weed, R. A.; Carlson, L. A.; Anderson, W. K.
A three-dimensional transonic-wing design algorithm for vector computers is developed, and the results of sample computations are presented graphically. The method incorporates the direct/inverse scheme of Carlson (1975), a Cartesian grid system with boundary conditions applied at a mean plane, and a potential-flow solver based on the conservative form of the full potential equation and using the ZEBRA II vectorizable solution algorithm of South et al. (1980). The accuracy and consistency of the method with regard to direct and inverse analysis and trailing-edge closure are verified in the test computations.
Craig, J. E.
A holographic interferometer system was installed in a 2X2 foot transonic wind tunnel. The system incorporates a modern, 10 pps, Nd:YAG pulsed laser which provides reliable operation and is easy to align. The spatial filtering requirements of the unstable resonator beam are described as well as the integration of the system into the existing Schieren system. A two plate holographic interferometer is used to reconstruct flow field data. For static wind tunnel models the single exposure holograms are recorded in the usual manner; however, for dynamic models such as oscillating airfoils, synchronous laser hologram recording is used.
Harris, Charles D.; Brooks, Cuyler W., Jr.
Modifications to the NASA Langley 8 Foot Transonic Pressure Tunnel in support of the Lamina Flow Control (LFC) Experiment included the installation of a honeymoon and five screens in the settling chamber upstream of the test section 41-long test section liner that extended from the upstream end of the test section contraction region, through the best section, and into the diffuser. The honeycomb and screens were installed as permanent additions to the facility, and the liner was a temporary addition to be removed at the conclusion of the LFC Experiment. These modifications are briefly described.
Adcock, J. B.
Real gas solutions for one-dimensional isentropic and normal-shock flows of nitrogen were obtained for a wide range of temperatures and pressures. These calculations are compared to ideal gas solutions and are presented in tables. For temperatures (300 K and below) and pressures (1 to 10 atm) that cover those anticipated for transonic cryogenic tunnels, the solutions are analyzed to obtain indications of the magnitude of inviscid flow simulation errors. For these ranges, the maximum deviation of the various isentropic and normal shock parameters from the ideal values is about 1 percent or less, and for most wind tunnel investigations this deviation would be insignificant.
Gentry, C. L., Jr.; Igoe, W. B.; Fuller, D. E.
The National Transonic Facility (NTF) is a pressurized cryogenic wind tunnel with a 2.5 m square test section. A 0.186-scale model of the NTF was used to simulate the aerodynamic performance of the components of the high-speed duct of the NTF. These components consist of a wide-angle diffuser, settling chamber, contraction section, test section, model support section, and high-speed diffuser. The geometry of the model tunnel, referred to as the diffuser flow apparatus is described, and some of its operating characteristics are presented.
This paper compares the results of pressure measurements made on the same test article with the same test matrix in three transonic wind tunnels. A comparison is presented of the unexplained variance associated with polar replicates acquired in each tunnel. The impact of a significance component of systematic (not random) unexplained variance is reviewed, and the results of analyses of variance are presented to assess the degree of significant systematic error in these representative wind tunnel tests. Total uncertainty estimates are reported for 140 samples of pressure data, quantifying the effects of within-polar random errors and between-polar systematic bias errors.
Doggett, Robert V., Jr.; Soistmann, David L.; Spain, Charles V.; Parker, Ellen C.; Silva, Walter A.
Transonic flutter boundaries are presented for two simple, 72 deg. sweep, low-aspect-ratio wing models. One model was an aspect-ratio 0.65 delta wing; the other model was an aspect-ratio 0.54 clipped-delta wing. Flutter boundaries for the delta wing are presented for the Mach number range of 0.56 to 1.22. Flutter boundaries for the clipped-delta wing are presented for the Mach number range of 0.72 to 0.95. Selected vibration characteristics of the models are also presented.
Allison, Dennis O.; Waggoner, E. G.
Computational predictions of the effects of wing contour modifications on maximum lift and transonic performance were made and verified against low speed and transonic wind tunnel data. This effort was part of a program to improve the maneuvering capability of the EA-6B electronics countermeasures aircraft, which evolved from the A-6 attack aircraft. The predictions were based on results from three computer codes which all include viscous effects: MCARF, a 2-D subsonic panel code; TAWFIVE, a transonic full potential code; and WBPPW, a transonic small disturbance potential flow code. The modifications were previously designed with the aid of these and other codes. The wing modifications consists of contour changes to the leading edge slats and trailing edge flaps and were designed for increased maximum lift with minimum effect on transonic performance. The prediction of the effects of the modifications are presented, with emphasis on verification through comparisons with wind tunnel data from the National Transonic Facility. Attention is focused on increments in low speed maximum lift and increments in transonic lift, pitching moment, and drag resulting from the contour modifications.
Luoma, Arvo A.
The longitudinal stability and control characteristics of a 1/30-scale model of the Republic XF-103 airplane were investigated in the Langley 8-foot transonic tunnel. The effect of speed brakes located at the end of the fuselage was also investigated. The main part of the investigation was made with internal flow in the model, but some data were obtained with no internal flow. The longitudinal stability and control at transonic-speeds appeared satisfactory. The transonic drag rise was small. The speed brakes had no adverse effects on longitudinal stability.
Seidel, D. A.; Batina, J. T.
The development, use and operation of the XTRAN2L program that solves the two dimensional unsteady transonic small disturbance potential equation are described. The XTRAN2L program is used to calculate steady and unsteady transonic flow fields about airfoils and is capable of performing self contained transonic flutter calculations. Operation of the XTRAN2L code is described, and tables defining all input variables, including default values, are presented. Sample cases that use various program options are shown to illustrate operation of XTRAN2L. Computer listings containing input and selected output are included as an aid to the user.
Mcdevitt, John B
The effects of one type of camber on the aerodynamic characteristics of rectangular wings at high subsonic and transonic speeds have been studied by applying the transonic similarity rules to the correlation of experimental data for a series of 18 cambered wings having NACA 63A2XX and 63A4XX sections, aspect ratios from 1 to 4, and thicknesses from 4 to 8 percent. The data were obtained by use of a transonic bump over a Mach number range of 0.6 to 1.1.
Kirtley, K. R.; Beach, T. A.; Rogo, Cass
A numerical simulation of a transonic mixed flow turbine stage has been carried out using an average passage Navier-Stokes analysis. The mixed flow turbine stage considered here consists of a transonic nozzle vane and a highly loaded rotor. The simulation was run at the design pressure ratio and is assessed by comparing results with those of an established throughflow design system. The three-dimensional aerodynamic loads are studied as well as the development and migration of secondary flows and their contribution to the total pressure loss. The numerical results indicate that strong passage vortices develop in the nozzle vane, mix out quickly, and have little impact on the rotor flow. The rotor is highly loaded near the leading edge. Within the rotor passage, strong spanwise flows and other secondary flows exist along with the tip leakage vortex. The rotor exit loss distribution is similar in character to that found in radial inflow turbines. The secondary flows and non-uniform work extraction also tend to significantly redistribute a non-uniform inlet total temperature profile by the exit of the stage.
Meyers, J. F.; Hunter, W. W., Jr.; Reubush, D. E.; Nichols, C. E., Jr.; Hepner, T. E.; Lee, J. W.
An investigation in the Langley 16-Foot Transonic Tunnel has been conducted in which a laser velocimeter was used to measure free-stream velocities from Mach 0.1 to 1.0 and the flow velocities along the stagnating streamline of a hemisphere-cylinder model at Mach 0.8 and 1.0. The flow velocity was also measured at Mach 1.0 along the line 0.533 model diameters below the model. These tests determined the performance characteristics of the dedicated two-component laser velocimeter at flow velocities up to Mach 1.0 and the effects of the wind tunnel environment on the particle-generating system and on the resulting size of the generated particles. To determine these characteristics, the measured particle velocities along the stagnating streamline at the two Mach numbers were compared with the theoretically predicted gas and particle velocities calculated using a transonic potential flow method. Through this comparison the mean detectable particle size (2.1 micron) along with the standard deviation of the detectable particles (0.76 micron) was determined; thus the performance characteristics of the laser velocimeter were established.
Keener, E. R.
Experimental surface pressure distributions and oil flow photographs are presented for a 0.90 m semispan model of NASA/Lockheed Wing C, a generic transonic, supercritical, low aspect ratio, highly 3-dimensional configuration. This wing was tested at the design angle of attack of 5 deg over a Mach number range from 0.25 to 0.96, and a Reynolds number range from 3.4 x 1,000,000 to 10 x 1,000,000. Pressures were measured with both the tunnel floor and ceiling suction slots open for most of the tests but taped closed for some tests to simulate solid walls. A comparison is made with the measured pressures from a small model in high Reynolds number facility and with predicted pressures using two three dimesional, transonic full potential flow wing codes: design code FLO22 (nonconservative) and TWING code (conservative). At the given design condition, a small region of flow separation occurred. At a Mach number of 0.82 the flow was unseparated and the surface flow angles were less than 10 deg, indicating that the boundary layer flow was not 3-D. Evidence indicate that wings that are optimized for mild shock waves and mild pressure recovery gradients generally have small 3-D boundary layer flow at design conditions for unseparated flow.
Statnikov, V.; Glatzer, C.; Meiß, J.-H.; Meinke, M.; Schröder, W.
Numerical simulations of the near wake of generic rocket configurations are performed at transonic and supersonic freestream conditions to improve the understanding of the highly intricate near wake structures. The Reynolds number in both flow regimes is 106 based on the main body diameter, i. e., specific freestream conditions of ESA's Ariane launcher trajectory. The geometry matches models used in experiments in the framework of the German Transregional Collaborative Research Center TRR40. Both axisymmetric wind tunnel models possess cylindrical sting supports, representing a nozzle to allow investigations of a less disturbed wake flow. A zonal approach consisting of a Reynolds averaged Navier-Stokes (RANS) and a large-eddy simulation (LES) is applied. It is shown that the highly unsteady transonic wake flow at Ma∞ = 0.7 is characterized by the expanding separated shear layer, while the Mach 6.0 wake is defined by a shock, expansion waves, and a recompression region. In both cases, an instantaneous view on the base characteristics reveals complex azimuthal flow structures even for axisymmetric geometries. The flow regimes are discussed by comparing the aerodynamic characteristics, such as the size of the recirculation region and the turbulent kinetic energy.
Menzies, Margaret A.; Kandil, Osama A.
The unsteady, three-dimensional, full Navier-Stokes (NS) equations and the Euler equations of rigid-body dynamics are sequentially solved to simulate the natural rolling response of slender delta wings of zero thickness at moderate to high angles of attack, to transonic and subsonic flows. The governing equations of fluid flow and dynamics of the present multi-disciplinary problem are solved using the time-accurate solution of the NS equations with the implicit, upwind, Roe flux-difference splitting, finite-volume scheme and a four-stage Runge-Kutta scheme, respectively. The main focus is to analyze the effect of Mach number and angle of attack on the leading edge vortices and their breakdown, the resultant rolling motion, and overall aerodynamic response of the wing. Three cases demonstrate the natural response of a 65 deg swept, cropped delta wing in a transonic flow with breakdown of the leading edge vortices and an 80 deg swept delta wing in a subsonic flow undergoing either damped or self-excited limit-cycle rolling oscillations as a function of angle of attack. Comparisons with an experimental investigation completes this study, validating the analysis and illustrating the complex details afforded by computational investigations.
Giel, P. W.; Thurman, D. R.; Lopez, I.; Boyle, R. J.; VanFossen, G. J.; Jett, T. A.; Camperchioli, W. P.; La, H.
Three-dimensional flow field measurements are presented for a large scale transonic turbine blade cascade. Flow field total pressures and pitch and yaw flow angles were measured at an inlet Reynolds number of 1.0 x 10(exp 6) and at an isentropic exit Mach number of 1.3 in a low turbulence environment. Flow field data was obtained on five pitchwise/spanwise measurement planes, two upstream and three downstream of the cascade, each covering three blade pitches. Three-hole boundary layer probes and five-hole pitch/yaw probes were used to obtain data at over 1200 locations in each of the measurement planes. Blade and endwall static pressures were also measured at an inlet Reynolds number of 0.5 x 10(exp 6) and at an isentropic exit Mach number of 1.0. Tests were conducted in a linear cascade at the NASA Lewis Transonic Turbine Blade Cascade Facility. The test article was a turbine rotor with 136 deg of turning and an axial chord of 12.7 cm. The flow field in the cascade is highly three-dimensional as a result of thick boundary layers at the test section inlet and because of the high degree of flow turning. The large scale allowed for very detailed measurements of both flow field and surface phenomena. The intent of the work is to provide benchmark quality data for CFD code and model verification.
Hu, Hong; Kandil, Osama A.
A hybrid boundary element finite volume method for unsteady transonic flow computation has been developed. In this method, the unsteady Euler equations in a moving frame of reference are solved in a small embedded domain (inner domain) around the airfoil using an implicit finite volume scheme. The unsteady full-potential equation, written in the same frame of reference and in the form of the Poisson equation. is solved in the outer domain using the integral equation boundary element method to provide the boundary conditions for the inner Euler domain. The solution procedure is a time-accurate stepping procedure, where the outer boundary conditions for the inner domain are updated using the integral equation -- boundary element solution over the outer domain. The method is applied to unsteady transonic flows around the NACA0012 airfoil undergoing pitching oscillation and ramp motion. The results are compared with those of an implicit Euler equation solver, which is used throughout a large computational domain, and experimental data.
Fry, K. A.; Bryanston-Cross, P.
The diagnostic use of quantitative laser flow visualization techniques has increased rapidly over recent years. The limitations imposed by conventional single point techniques such as laser Doppler anemometry are addressed and how they have been overcome by the development of a new family of whole field measurement techniques is demonstrated. In particular near instantaneous whole field velocity data was obtained in a relatively hostile, industrial 2.74 m x 2.44 m transonic wind tunnel (TWT) at the Aircraft Research Association (ARA). The techniques were evaluated for their suitability for making quantitative measurements in the wing/pylon region of a model wing and engine combination. Three optical diagnostic techniques were successfully developed within the context of the ARA facility. The first technique, laser light sheet (LLS), combines the operation of a pulse laser and video capture system to provide a 'real time' visualization of the flow, whereas a second pulse laser technique, Particle Image Velocimetry (PIV) can be used to make specific quantitative whole field instantaneous velocity measurements. The third method, holography, was used to produce a stored three dimensional visualization of the unsteady and shock wave features of the transonic flow in the gully region. A description is made of their installation and operation, and examples are presented of current test results.
Keith, J. S.; Ferguson, D. R.; Merkle, C. L.; Heck, P. H.; Lahti, D. J.
The formulation and development of a computer analysis for the calculation of streamlines and pressure distributions around two-dimensional (planar and axisymmetric) isolated nacelles at transonic speeds are described. The computerized flow field analysis is designed to predict the transonic flow around long and short high-bypass-ratio fan duct nacelles with inlet flows and with exhaust flows having appropriate aerothermodynamic properties. The flow field boundaries are located as far upstream and downstream as necessary to obtain minimum disturbances at the boundary. The far-field lateral flow field boundary is analytically defined to exactly represent free-flight conditions or solid wind tunnel wall effects. The inviscid solution technique is based on a Streamtube Curvature Analysis. The computer program utilizes an automatic grid refinement procedure and solves the flow field equations with a matrix relaxation technique. The boundary layer displacement effects and the onset of turbulent separation are included, based on the compressible turbulent boundary layer solution method of Stratford and Beavers and on the turbulent separation prediction method of Stratford.
Mirowitz, L. I.
Transonic flight flutter tests of the XF3H-1 Demon Airplane were conducted utilizing a frequency response technique in which the oscillating rudder provides the means of system excitation. These tests were conducted as a result of a rudder flutter incident in the transonic speed range. The technique employed is presented including a brief theoretical development of basic concepts. Test data obtained during the flight are included and the method of interpretation of these data is indicated. This method is based on an impedance matching technique. It is shown that an artificial stabilizing device, such as a damper, may be incorporated in the system for test purposes without complicating the interpretation of the test results of the normal configuration. Data are presented which define the margin of stability introduced to the originally unstable rudder by design changes which involve higher control system stiffness and external damper. It is concluded that this technique of flight flutter testing is a feasible means of obtaining flutter stability information in flight.
Kodzwa, Paul M.; Eaton, John K.
This paper presents steady-state recovery temperature and heat transfer coefficient measurements on the pressure surface of a modern, highly cambered transonic airfoil. These measurements were collected with a peak Mach number of 1.5 and a maximum turbulence intensity of 30%. We used a single passage model to simulate the idealized two-dimensional flow path between rotor blades in a modern transonic turbine. This set up offered a simpler construction than a linear cascade, yet produced an equivalent flow condition. We performed validated high accuracy (±0.2°C) surface temperature measurements using wide-band thermochromic liquid crystals allowing separate measurements of the previously listed parameters with the same heat transfer surface. We achieved maximum heat transfer coefficient uncertainties that were equivalent to similar investigations (±10%). Two key observations are the heat transfer coefficient along the aft portion of the airfoil is sensitive to the surface heat flux and is highly insensitive to the level of freestream turbulence. Possible explanations for these observations are discussed.
Al-Saadi, Jassim A.
A computational simulation of a transonic wind tunnel test section with longitudinally slotted walls is developed and described herein. The nonlinear slot model includes dynamic pressure effects and a plenum pressure constraint, and each slot is treated individually. The solution is performed using a finite-difference method that solves an extended transonic small disturbance equation. The walls serve as the outer boundary conditions in the relaxation technique, and an interaction procedure is used at the slotted walls. Measured boundary pressures are not required to establish the wall conditions but are currently used to assess the accuracy of the simulation. This method can also calculate a free-air solution as well as solutions that employ the classical homogeneous wall conditions. The simulation is used to examine two commercial transport aircraft models at a supercritical Mach number for zero-lift and cruise conditions. Good agreement between measured and calculated wall pressures is obtained for the model geometries and flow conditions examined herein. Some localized disagreement is noted, which is attributed to improper simulation of viscous effects in the slots.
The numerical simulation of 3-D transonic flow about a system of propeller blades is investigated. In particular, it is shown that the use of helical coordinates significantly simplifies the form of the governing equation when the propeller system is assumed to be surrounded by an irrotational flow field of an inviscid fluid. The unsteady small disturbance equation, valid for lightly loaded blades and expressed in helical coordinates, is derived from the general blade-fixed potential equation, given for an arbitrary coordinate system. The use of a coordinate system which inherently adapts to the mean flow results in a disturbance equation requiring relatively few terms to accurately model the physics of the flow. Furthermore, the helical coordinate system presented here is novel in that it is periodic in the circumferential direction while, simultaneously, maintaining orthogonal properties at the mean blade locations. The periodic characteristic allows a complete cascade of blades to be treated, and the orthogonality property affords straightforward treatment of blade boundary conditions. An ADI numerical scheme is used to compute the solution of the steady flow as an asymptotic limit of an unsteady flow. As an example of the method, solutions are presented for subsonic and transonic flow about a 5 percent thick bicircular arc blade of an 8-bladed cascade. Both high and low advance ratio cases are computed and include a lifting as well as nonlifting cases. The nonlifting solutions obtained are compared to solutions from a Euler code.
Hah, Chunill; Rabe, Douglas; Scribben, Angie
In the present study, unsteady flow interaction between an ultra-compact inlet and a transonic fan stage is investigated. Future combat aircraft require ultra-compact inlet ducts as part of an integrated, advanced propulsion system to improve air vehicle capability and effectiveness to meet future mission needs. The main purpose of the study is to advance the current understanding of the flow interaction between two different ultra-compact inlets and a transonic fan for future design applications. Both URANS and LES approaches are used to calculate the unsteady flow field and are compared with the available measured data. The present study indicates that stall inception is mildly affected by the distortion pattern generated by the inlet with the current test set-up. The numerical study indicates that the inlet distortion pattern decays significantly before it reaches the fan face for the current configuration. Numerical results with a shorter distance between the inlet and fan show that counter-rotating vortices near the rotor tip due to the serpentine diffuser affects fan characteristics significantly.
Suzuki, Kojiro; Abe, Takashi
The boundary-layer transition on the EXPRESS reentry capsule at transonic and supersonic speeds is studied experimentally by the wind tunnel tests. For the diagnostic of the turbulent transition of the boundary layer, the China-clay method is used. The experimental results clarify that when the freestream Mach number increases, the transition point moves downstream on the body surface and the distance between the beginning of the transition and its completion to the fully turbulent flow becomes larger. The effects of the freestream Mach number on the location of the boundary-layer transition are described successfully in terms of two nondimensional quantities, that is, the transition Reynolds number and the local Mach number at the boundary-layer edge. The oil-flow pictures reveal that in the transonic regime, the separation bubble is formed at the junction between the blunt nose and the conical part of the body and therefore the transition begins behind the reattachment point of the separation bubble. The effects of the turbulent transition on the aerodynamic characteristics of the reentry body are investigated by using the technique of the boundary-layer trip and the experimental results show that the aerodynamic characteristics of the EXPRESS reentry vehicle are not sensitive to the boundary-layer transition.
Zuniga, Fanny A.; Drake, Aaron; Kennelly, Robert A., Jr.; Koga, Dennis J.; Westphal, Russell V.
A flight experiment, conducted at NASA Dryden Flight Research Center, investigated the effects of surface excrescences, specifically gaps and steps, on boundary-layer transition in the vicinity of a leading edge at transonic flight conditions. A natural laminar flow leading-edge model was designed for this experiment with a spanwise slot manufactured into the leading-edge model to simulate gaps and steps like those present at skin joints of small transonic aircraft wings. The leading-edge model was flown with the flight test fixture, a low-aspect ratio fin mounted beneath an F-104G aircraft. Test points were obtained over a unit Reynolds number range of 1.5 to 2.5 million/ft and a Mach number range of 0.5 to 0.8. Results for a smooth surface showed that laminar flow extended to approximately 12 in. behind the leading edge at Mach number 0.7 over a unit Reynolds number range of 1.5 to 2.0 million/ft. The maximum size of the gap-and-step configuration over which laminar flow was maintained consisted of two 0.06-in. gaps with a 0.02-in. step at a unit Reynolds number of 1.5 million/ft.
Kulfan, R. M.; Neumann, F. D.; Nisbet, J. W.; Mulally, A. R.; Murakami, J. K.; Noble, E. C.; Mcbarron, J. P.; Stalter, J. L.; Gimmestad, D. W.; Sussman, M. B.
An initial design study of high-transonic-speed transport aircraft has been completed. Five different design concepts were developed. These included fixed swept wing, variable-sweep wing, delta wing, double-fuselage yawed-wing, and single-fuselage yawed-wing aircraft. The boomless supersonic design objectives of range=5560 Km (3000 nmi), payload-18 143 kg (40 000lb), Mach=1.2, and FAR Part 36 aircraft noise levels were achieved by the single-fuselage yawed-wing configuration with a gross weight of 211 828 Kg (467 000 lb). A noise level of 15 EPNdB below FAR Part 36 requirements was obtained with a gross weight increase to 226 796 Kg (500 000 lb). Although wing aeroelastic divergence was a primary design consideration for the yawed-wing concepts, the graphite-epoxy wings of this study were designed by critical gust and maneuver loads rather than by divergence requirements. The transonic nacelle drag is shown to be very sensitive to the nacelle installation. A six-degree-of-freedom dynamic stability analysis indicated that the control coordination and stability augmentation system would require more development than for a symmetrical airplane but is entirely feasible. A three-phase development plan is recommended to establish the full potential of the yawed-wing concept.
Schuster, David M.
Tests conducted in the NASA Langley Research Center Transonic Dynamics Tunnel (TDT) assess the aerodynamic characteristics of a splitter plate used to test some semispan models in this facility. Aerodynamic data are analyzed to determine the effect of the splitter plate on the operating characteristics of the TDT, as well as to define the range of conditions over which the plate can be reasonably used to obtain aerodynamic data. Static pressures measurements on the splitter plate surface and the equipment fairing between the wind tunnel wall and the splitter plate are evaluated to determine the flow quality around the apparatus over a range of operating conditions. Boundary layer rake data acquired near the plate surface define the viscous characteristics of the flow over the plate. Data were acquired over a range of subsonic, transonic and supersonic conditions at dynamic pressures typical for models tested on this apparatus. Data from this investigation should be used as a guide for the design of TDT models and tests using the splitter plate, as well as to guide future splitter plate design for this facility.
Piatak, David J.; Sekula, Martin K.; Rausch, Russ D.; Florance, James R.; Ivanco, Thomas G.
Fluctuating aerodynamic loads are a significant concern for the structural design of a launch vehicle, particularly while traversing the transonic flight environment. At these trajectory conditions, unsteady aerodynamic pressures can excite the vehicle dynamic modes of vibration and result in high structural bending moments and vibratory environments. To ensure that vehicle structural components and subsystems possess adequate strength, stress, and fatigue margins in the presence of buffet and other environments, buffet forcing functions are required to conduct the coupled load analysis of the launch vehicle. The accepted method to obtain these buffet forcing functions is to perform wind-tunnel testing of a rigid model that is heavily instrumented with unsteady pressure transducers designed to measure the buffet environment within the desired frequency range. Two wind-tunnel tests of a 3 percent scale rigid buffet model have been conducted at the Langley Research Center Transonic Dynamics Tunnel (TDT) as part of the Space Launch System (SLS) buffet test program. The SLS buffet models have been instrumented with as many as 472 unsteady pressure transducers to resolve the buffet forcing functions of this multi-body configuration through integration of the individual pressure time histories. This paper will discuss test program development, instrumentation, data acquisition, test implementation, data analysis techniques, and several methods explored to mitigate high buffet environment encountered during the test program. Preliminary buffet environments will be presented and compared using normalized sectional buffet forcing function root-meansquared levels along the vehicle centerline.