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A procedure for tailoring a blended wing-body configuration to reduce its computed wave drag is described. The method utilizes an iterative algorithm within the framework of first-order linear theory. Four computed examples are included. In each case, the zero-lift wave drag was reduced without an increase in the drag due to lift.
Simulation of high-speed three-dimensional flow about blended wing-body combinations is investigated. A finite-volume explicit scheme with Runge-Kutta time integration is used to solve the compressible Euler equations in order to simulate the flow. The method, augmented by carefully chosen dissipative terms and convergence accelerators such as enthalpy damping and maximum local time-stepping, has been found to be very efficient in solving high-speed flows involving strong shocks. An analytic method for generating the body geometry and an algebraic method for quasi-three-dimensional grid generation are described. Results are presented for various blended wing-body configurations at different Mach numbers and angles of attack.
Force balance and wing pressure data were obtained on a 0.017-Scale Model of a blended-wing-body configuration (without a simulated propulsion system installation) to validate the capability of computational fluid dynamic codes to predict the performance of such thick sectioned subsonic transport configurations. The tests were conducted in the National Transonic Facility of the Langley Research Center at Reynolds numbers from 3.5 to 25.0 million at Mach numbers from 0.25 to 0.86. Data were obtained in the pitch plane only at angles of attack from -1 to 8 deg at Mach numbers greater than 0.25. A configuration with winglets was tested at a Reynolds number of 25.0 million at Mach numbers from 0.83 to 0.86.
Methods for prediction of centerline shock layer thickness and pressure distribution on delta wing body configurations
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.
A direct algebraic method was developed and applied to generate three dimensional grids around wing-body configurations. The method used is a generalized transfinite interpolation method which generates the desired coordinate transformation using geometric data only on the boundaries of the domain of interest. The geometric data that can be specified includes not only coordinates on the boundaries but also out-of-surface parametric derivatives that give a very precise control over the transformation in the vicinity of the surface. In addition to this, the method gives good control over the stretching of the mesh between different boundaries.
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Unsteady Navier-Stokes computations are conducted for transonic flows over a wing-body configuration undergoing prescribed ramp motions. The ramp motion from 0 to 15 deg includes angles of attack where vortex breakdown is observed experimentally. The vortex breakdown is found to be delayed until after the ramp motion ends. The dynamic effects on the loads are also demonstrated. The moment coefficient is found to be sensitive to the effect of the virtual mass of the fluid. To verify the numerical results, a grid refinement study has been performed for both steady and unsteady flow conditions, by using up to approximately one million grid points. The results show that the coarse grid can give integrated quantities reasonably well, whereas the finer grids give a more detailed flow structure. Comparisons are also made with available steady-state experimental data.
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A wind tunnel study at Mach 0.4 was conducted for a slender wing-body configuration with a leading edge vortex flap of curved planform that is deflectable about a 74 degree swept hinge line. The basic data consist of a unique combination of longitudinal aerodynamic, surface pressure, and vortex flap hinge-moment measurements on a common model. The longitudinal aerodynamic, pressure and hinge-moment data are presented without analysis in tabular format. Plots of the tabulated pressure data are also given.
An experimental investigation has been conducted to determine the effect of wing leading edge sweep and wing translation on the aerodynamic characteristics of a wing body configuration at a free stream Mach number of about 6 and Reynolds number (based on body length) of 17.9 x 10 to the 6th power. Seven wings with leading edge sweep angles from -20 deg to 60 deg were tested on a common body over an angle of attack range from -12 deg to 10 deg. All wings had a common span, aspect ratio, taper ratio, planform area, and thickness ratio. Wings were translated longitudinally on the body to make tests possible with the total and exposed mean aerodynamic chords located at a fixed body station. Aerodynamic forces were found to be independent of wing sweep and translation, and pitching moments were constant when the exposed wing mean aerodynamic chord was located at a fixed body station. Thus, the Hypersonic Isolation Principle was verified. Theory applied with tangent wedge pressures on the wing and tangent cone pressures on the body provided excellent predictions of aerodynamic force coefficients but poor estimates of moment coefficients.
An exploratory, experimental, and theoretical investigation was made of a cambered, twisted, and blended wing-body concept with and without integral canard surfaces. Theoretical calculations of the static longitudinal and lateral aerodynamic characteristics of the wing-body configurations were compared with the characteristics obtained from tests of a model in the Langley Unitary Plan wind tunnel. Mach numbers of 1.5, 1.8, and 2.0 and a Reynolds number per meter of 6.56 million were used in the calculations and tests. Overall results suggest that planform selection is extremely important and that the supplemental application of new calculation techniques should provide a process for the design of supersonic wings in which spanwise distribution of upwash and leading-edge thrust might be rationally controlled and exploited.
The thin-layer Navier-Stokes equations are solved for the flow about a coplanar close-coupled canard-wing-body configuration at a transonic Mach number of 0.90 and at angles of attack ranging from 0 to 12 degrees. The influence of the canard on the wing flowfield, including canard-wing vortex interaction and wing vortex breakdown, is investigated. A study of canard downwash and canard leading-edge vortex effects, which are the primary mechanisms of the canard-wing interaction, is emphasized. Comparisons between the computations and experimental measurements of surface pressure coefficients, lift, drag and pitching moment data are favorable. A grid refinement study for configurations with and without canard shows that accurate results are obtained using a refined grid for angles of attack where vortex burst is present. At an angle of attack of approximately 12 deg, favorable canard-wing interaction which delays wing vortex breakdown is indicated by the computations and is in good agreement with experimental findings.
An investigation was conducted in the Langley low turbulence pressure tunnel to determine the effects of wing leading edge radius and Reynolds number on the longitudinal aerodynamic characteristics of a series of highly swept wing-body configurations. The tests were conducted at Mach numbers below 0.30, angles of attack up to 16 deg, and Reynolds numbers per meter from 6.57 million to 43.27 million. The wings under study in this investigation had leading edge sweep angles of 61.7 deg, 64.61 deg, and 67.01 deg in combination with trailing edge sweep angles of 0 deg and 40.6 deg. The leading edge radii of each wing planform could be varied from sharp to nearly round.
An investigation has been conducted in the Langley 1/3-meter transonic cryogenic tunnel to determine the effects of varying Reynolds number on the boattail drag of wing-body configurations at subsonic speeds. Two boattailed cone-cylinder nacelle models were tested with a 60 deg delta wing at an angle of attack of 0 deg. Reynolds number, based on model length, was varied from about 2.5 million to 67 million. Even though the presence of the wing had large effects on the boattail pressure coefficients, the results of this investigation were similar to those previously found for a series of isolated boattails. Boattail pressure coefficients in the expansion region became more negative with increasing Reynolds number, while those in the recompression region became more positive. These two effects were compensating, and as a result, there was virtually no effect of Reynolds number on boattail pressure drag.
A hybrid transition trip-dot sizing and placement test technique was developed in support of recent experimental research on a hybrid wing-body configuration under study for the NASA Environmentally Responsible Aviation project. The approach combines traditional methods with Computational Fluid Dynamics. The application had three-dimensional boundary layers that were simulated with either fully turbulent or transitional flow models using established Reynolds-Averaged Navier-Stokes methods. Trip strip effectiveness was verified experimentally using infrared thermography during a low-speed wind tunnel test. Although the work was performed on one specific configuration, the process was based on fundamental flow physics and could be applicable to other configurations.
The thin-layer Navier-Stokes equations are solved numerically to investigate the effects of canard vertical position on a close-coupled canard-wing-body configuration at a transonic Mach number of 0.90 and angles of attack ranging from -2 to 12 degrees. Canard-wing interactions are investigated for high-, mid- and low-canard positions. The computational results show favorable canard-wing interactions for the high- and mid-canard configurations. The unfavorable lift and drag characteristics for the low-canard configuration are examined by analyses of the low-canard flowfield structure.
An investigation was conducted in a transonic cryogenic tunnel to determine the effect of varying Reynolds number on the boattail drag of several wing-body configurations. This study was made at 0 deg angle of attack at Mach numbers from 0.6 to 0.9 for Reynolds numbers up to 67 x 1 million (based on distance from the nose to the start of the boattail). Results indicate that as the Reynolds number was increased the boattail static pressure coefficients in the expansion region of the boattail became more negative while those in the recompression region became more positive. Results show that there was only a small effect of Reynolds number of boattail pressure drag.