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Application of low drag-due-to-lift concept of linearized theory to the studies of twist and camber of isolated wing and wing-body configurations of supersonic aircraft
Four student design teams produced conceptual designs for a research vehicle to develop the supersonic combustion ramjet (scramjet) engines necessary for efficient hypersonic flight. This research aircraft would provide flight test data for prototype scramjets that is not available in groundbased test facilities. The design specifications call for a research aircraft to be launched from a carrier aircraft at 40,000 feet and a Mach number of 0.8. The aircraft must accelerate to Mach 6 while climbing to a 100,000 foot altitude and then ignite the experimental scramjet engines for acceleration to Mach 10. The research vehicle must then be recovered for another flight. The students responded with four different designs, two piloted waverider configurations, and two unmanned vehicles, one with a blended body-wing configuration, the other with a delta wing shape. All aircraft made use of an engine database provided by the General Electric Aircraft Engine Group; both turbofan ramjet and scramjet engine performance using liquid hydrogen fuel was available. Explained here are the students' conceptual designs and the aerodynamic and propulsion concepts that made their designs feasible.
A flight investigation has been made to determine the longitudinal stability and control characteristics of a 60 0 delta-wing-canard missile configuration with an exposed wing-canard area ratio of 16:1. The results presented include the longitudinal stability derivatives, control effectiveness, and drag characteristics for a Mach number range of 0.75 to 1.80 and are compared with the results of a similar configuration having larger 6ontrols. Stability characteristics are also presented from the flights of an interdigitated canard configuration at a Mach number of 2.08 and a wing-body configuration at Mach numbers of 1.25 to 1.45. The stability derivatives varied gradually with Mach number with the exception of the damping-in-pitch derivative. Aerodynamic damping in pitch decreased to a minimum at a Mach number of 1.0 3, then increased to a peak value at a Mach number of 1.26 followed by a gradual decrease at higher Mach numbers. The aerodynamic-center location of the in-line canard configuration shifted rearward 13 percent of the mean aerodynamic chord at transonic speeds. The pitching-moment curve slope was 25 percent greater for the model having no canards than for the in-line configuration. No large effects of interdigitation were noted in the stability derivatives. Pitching effectiveness of the in-line configuration was maintained throughout the Mach number range. A comparison of the stability and control characteristics of two canard configurations having different area controls showed that decreasing the control area 44 percent decreased the pitching effectiveness proportionally, shifted the aerodynamic-center location rearward 9 to 14 percent of the mean aerodynamic chord, and reduced the total hinge moments required for 10 trimmed flight about 50 percent at transonic speeds.
An overview is presented of 32 formal papers and 7 open session papers. Topics covered include: (1) studies of configurations of practical interest; (2) mathematical modelling and supporting investigations of slender wings, bodies of revolution, and body-wing configurations; (3) design methods; and (4) air intakes.
The wing-body aerodynamics data base consists of a series of CFD (Computational Fluid Dynamics) simulations about a generic wing body configuration consisting of a ogive-circular-cylinder fuselage and a simple symmetric wing mid-mounted on the fuselage. Solutions have been obtained for Nonlinear Potential (P), Euler (E) and Navier-Stokes (N) solvers over a range of subsonic and transonic Mach numbers and angles of attack. In addition, each solution has been computed on a series of grids, coarse, medium and fine to permit an assessment of grid refinement errors.
This report describes the development of a novel structural solution that addresses the demanding fuselage loading requirements for the Hybrid Wing or Blended Wing Body configurations that are described in NASA NRA subtopic A2A.3, "Materials and Structures for Wing Components and Non-Circular Fuselage." The phase I portion of this task includes a comprehensive finite element model-based structural sizing exercise performed using the BWB airplane configuration to generate internal loads and fuselage panel weights for an advanced Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) structural concept. An accompanying element-level test program is also described which substantiates the analytical results and calculation methods used in the trade study. The phase II plan for the continuation of this research is also included herein.
Configuration design at Ames was carried out with the SYN87-SB (single block) Euler code using a 193 x 49 x 65 C-H grid. The Euler solver is coupled to the constrained (NPSOL) and the unconstrained (QNMDIF) optimization packages. Since the single block grid is able to model only wing-body configurations, the nacelle/diverter effects were included in the optimization process by SYN87's option to superimpose the nacelle/diverter interference pressures on the wing. These interference pressures were calculated using the AIRPLANE code. AIRPLANE is an Euler solver that uses a unstructured tetrahedral mesh and is capable of computations about arbitrary complete configurations. In addition, the buoyancy effects of the nacelle/diverters were also included in the design process by imposing the pressure field obtained during the design process onto the triangulated surfaces of the nacelle/diverter mesh generated by AIRPLANE. The interference pressures and nacelle buoyancy effects are added to the final forces after each flow field calculation. Full details of the (recently enhanced) ghost nacelle capability are given in a related talk. The pseudo nacelle corrections were greatly improved during this design cycle. During the Ref H and Cycle 1 design activities, the nacelles were only translated and pitched. In the cycle 2 design effort the nacelles can translate vertically, and pitch to accommodate the changes in the lower surface geometry. The diverter heights (between their leading and trailing edges) were modified during design as the shape of the lower wing changed, with the drag of the diverter changing accordingly. Both adjoint and finite difference gradients were used during optimization. The adjoint-based gradients were found to give good direction in the design space for configurations near the starting point, but as the design approached a minimum, the finite difference gradients were found to be more accurate. Use of finite difference gradients was limited by the CPU time limit available on the Cray machines. A typical optimization run using finite difference gradients can use only 30 to 40 design variables and one optimization iteration within the 8 hour queue limit for the chosen grid size and convergence level. The efficiency afforded by the adjoint method allowed for 50-120 design variables and 5-10 optimization iterations in the 8 hour queue. Geometric perturbations to the wing and fuselage were made using the Hicks/Henne (HH) shape functions. The HH functions were distributed uniformly along the chords of the wing defining sections and lofted linearly. During single-surface design, constraints on thickness and volume at selected wing stations were imposed. Both fuselage camber and cross-sectional area distributions were permitted to change during design. The major disadvantage to the use of these functions is the inherent surface waviness produced by repeated use of such functions. Many smoothing operations were required following optimization runs to produce a configuration with reasonable smoothness. Wagner functions were also used on the wing sections but were never used on the fuselage. The Wagner functions are a family of increasingly oscillatory functions that have also been used extensively in airfoil design. The leading and trailing edge regions of the wing were designed by use of polynomial and monomial functions respectively. Twist was attempted but was abandoned because of little performance improvement available from changing the baseline twist.
Flat top wing body configuration effects on aerodynamic characteristics of supersonic aircraft
The NASA Environmentally Responsible Aviation (ERA) Project sponsored a series of computational and experimental investigations of the propulsion and airframe integration issues associated with Hybrid-Wing-Body (HWB) or Blended-Wing-Body (BWB) configurations. NASA collaborated with Boeing Research and Technology (BR&T) to conduct this research on a new twin-engine Boeing BWB transport configuration. The experimental investigations involved a series of wind tunnel tests with a 5.75-percent scale model conducted in two low-speed wind tunnels. This testing focused on the basic aerodynamics of the configuration and selection of the leading edge Krueger slat position for takeoff and landing. This paper reviews the results and analysis of these low-speed wind tunnel tests.
This paper evaluates the capability of the constant pressure panel method (CPM) code to predict unsteady aerodynamic pressures, lift and moment distributions, and generalized forces for general wing-body configurations in supersonic flow. Stability derivatives are computed and correlated for the X-29 and an Oblique Wing Research Aircraft, and a flutter analysis is carried out for a wing wind tunnel test example. Most results are shown to correlate well with test or published data. Although the emphasis of this paper is on evaluation, an improvement in the CPM code's handling of intersecting lifting surfaces is briefly discussed. An attractive feature of the CPM code is that it shares the basic data requirements and computational arrangements of the doublet lattice method. A unified code to predict unsteady subsonic or supersonic airloads is therefore possible.
A wind-tunnel investigation was made to determine the longitudinal- and lateral-stability derivatives of a flat-top wing-body configuration at Mach numbers from 0.22 to 0.90 and Reynolds numbers of 3.5 and 17 million. The wing had a leading-edge sweepback of 78.9 deg and a cathedral of 45 deg on the outer panels. The tests included the determination of the effectiveness of elevon and rudder controls and also an investigation of ground effects. The model was tested at angles of attack up to 28 deg and angles of sideslip up to 18 deg. The dynamic response of this configuration has been determined from the wind-tunnel data for a simulated airplane having a wing loading of 17.7 pounds per square foot. The longitudinal data show a forward shift in aerodynamic center of 10 percent of the mean aerodynamic chord as the lift coefficient is increased above 0.1. Although flown in the lift range of decreasing stability, the simulated airplane did not encounter pitch-up in maneuvers initiated from steady level flight with zero static margin unless a load factor of 2.2 was exceeded. This maneuver margin was provided by a large value of pitching moment due to pitching velocity. The number of cycles to damp the Dutch roll mode to half amplitude, the time constants of the roll subsidence and spiral divergence modes, and control effectiveness in roll are computed. The lateral stability is shown to be positive but is marginal in meeting the military specifications for today's aircraft. An analog computer study has been made in five degrees of freedom (constant velocity) which illustrates that the handling characteristics are satisfactory. Several programed rolling maneuvers and coordinated turns also illustrate the handling qualities of the airplane.
This paper deals with designing a thrust distribution strategy when a Turboelectric Distributed Propulsion (TeDP) system of 16 embedded propulsors is installed on an aerodynamically optimized hybrid wing-body configuration. This HWB previously designed to satisfy conditions of trim, longitudinally static stability and specific cargo space is employed as the baseline configuration for the current study of seeking an optimal propulsion/power system. According to the nature of the entrance flow condition for each distributed propulsion passage in hybrid wing-body aircraft, the ingested boundary layer thickness differs and results in different propulsive reaction. An optimal distribution of thrust and power output is determined by how the system utilizes the propulsive characteristics of each passage. The design space and the number of design variables are selected and described accordingly. An actuator disk model is employed to model thrust generation and shaft power from the propulsor. To carry out the optimization of the propulsion/power system on a computationally expensive CFD model, a Kriging method in conjunction with a Genetic Algorithm (GA) is applied. Throughout the design process, the propulsion performances of the sampled propulsion/power system are analyzed and compared to those of a clean flow engine. The performance metrics includes mass flow rate, fan pressure ratio besides the thrust and shaft power. Minimization of total shaft power from the distributed engine is performed at multiple thrust levels. The benefit of boundary layer ingestion propulsion system is quantified via comparison of thrust equivalent, shaft power and mass flow equivalent clean flow engines with CFD based system design.
An experimental study was performed at supersonic speeds to measure wing and body spanwise pressure distributions on an axisymmetric-body delta wing model on which the wing vertical location on the body was systematically varied from low- to high-mounted positions. In addition, for two of these positions both horizontal and radial wing angular orientations relative to the body were tested, and roll angle effects were investigated for one of the positions. Seven different wing-body configurations and a body-alone configuration were studied. The test was conducted at Mach numbers from 1.70 to 2.86 at angles of attack from about -4 deg to 24 deg. Pressure orifices were located at three longitudinal stations on each wing-body model, and at each station the orifices were located completely around the body, along the lower surface of the right wing (looking upstream), and along the upper surface of the left wing. All pressure coefficient data are tabulated and selected samples are shown graphically to illustrate the effects of the test variables. The effects of angle of attack, roll angle, Mach number, longitudinal station, wing vertical location, wing angular orientation, and wing-body juncture are analyzed. The vertical location of the wing on the body had a very strong effect on the body pressures. For a given angle of attack at a roll angle of 0 deg, the pressures were virtually constant in the spanwise direction across the windward surfaces of the wing-body combination. Pressure-relieving, channeling, and vortex effects were noted in the data.
The results of a study to design a High-Speed Civilian Transport (HSCT) using the blended wing-body configuration are presented. The HSCT is a Mach 2 to 5 transport aircraft designed to compete with current commercial aircraft. The subjects discussed are sizing, configuration, aerodynamics, stability and control, propulsion, performance, structures and pollution effects.
The influence of upper-surface nacelle exhaust flow on the aerodynamic characteristics of a supersonic cruise aircraft research configuration was investigated in a 16 foot transonic tunnel over a range of Mach numbers from 0.60 to 1.20. The arrow-wing transport configuration with engines suspended over the wing was tested at angles of attack from -4 deg to 6 deg and jet total pressure ratios from 1 to approximately 13. Wing-tip leading edge flap deflections of -10 deg to 10 deg were tested with the wing-body configuration. Various nacelle locations (chordwise, spanwise, and vertical) were tested over the ranges of Mach numbers, angles of attack, and jet total-pressure ratios. The results show that reflecting the wing-tip leading edge flap from 0 deg to -10 deg increased the maximum lift-drag ratio by 1.0 at subsonic speeds. Jet exhaust interference effects were negligible.
An experimental investigation was conducted in the Langley 16-Foot Transonic Tunnel at free-stream Mach numbers from 0.70 to 0.82 and angles of attack from -2.5 to 4.0 degrees to determine the integration effects of pylon-mounted underwing forward and rearward separate-flow, flow-through nacelles on a high-wing transonic transport configuration. The results showed that the installed drag of the nacelle/pylon in the rearward location was slightly less than that of the nacelle/pylon in the forward location. This reduction was due to the reduction in calculated skin friction of the nacelle/pylon configuration. In all cases the combined value of form, wave, and interference drag was excessively high. However, the configuration with the nacelle/pylon in a rearward location produced an increase in lift over that of the basic wing-body configuration.
Positions of the primary vortex flow reattachment line and longitudinal aerodynamic data were obtained at Mach number 0.3 for a systematic series of vortex flaps on delta wing body configurations with leading edge sweeps of 50, 58, 66, and 74 deg. The investigation was performed to study the parametric effects of wing sweep, vortex flap geometry and deflection, canards, and trailing edge flaps on the location of the primary vortex reattachment line relative to the flap hinge line. The vortex reattachment line was located via surface oil flow photographs taken at selected angles of attack. Force and moment measurements were taken over an angle of attack range of -1 deg to 22 deg at zero sideslip angle for many configurations to further establish the data base and to assess the aforementioned parametric effects on longitudinal aerodynamics. Both the flow reattachment and aerodynamic data are presented.
An aerodynamic optimization procedure developed to minimize the drag to lift ratio of an aircraft configuration: wing - body - tail, in accordance with engineering restrictions, is described. An algorithm developed to search a hypersurface with 18 dimensions, which define an aircraft configuration, is discussed. The results, when considered from the aerodynamic point of view, indicate the optimal configuration is one that combines a lifting fuselage with a canard.