Supersonic Transport Configuration Research Articles

Inverse Design Method for Low-Boom Supersonic Transport with Lift Constraint

The low-boom inverse design method is of great interest for the preliminary design of supersonic transport (SST), owing to its good capability of reducing sonic boom at low computational cost. However, the main challenge is how to prescribe an attainable target. This paper proposes an inverse design method using a new target-generation strategy to address this challenge. Unlike existing methods such as the Jones–Seebass–George–Darden method and the mixed-fidelity inverse design method [the Li–Shields–Geiselhart (LSG) method], the proposed method directly parameterizes the near-field overpressure distribution of a reference configuration and minimizes the ground perceived level in decibels (PLdB) predicted by solving the augmented Burgers equation to obtain an optimal target distribution, subject to cruise lift and volume constraints. The proposed method is demonstrated by the Japan Aerospace Exploration Agency (JAXA) Wing Body case, which shows that it mitigates sonic boom under lift constraint and achieves a reduction of 2.07 PLdB more than the LSG method. The proposed method is applied to the low-boom design of a large SST configuration called NPU-T7104, which is close in size to the Tupolev Tu-144; results show that aerodynamic performance is maintained while shock wave coalescence is effectively prevented, and ground-boom intensity is dramatically reduced.

Numerical Investigation of Influence of Mesh Property in Nearfield Sonic Boom Prediction

To investigate the influence of mesh property in nearfield sonic boom prediction, the over-pressures of several typical configurations are commutated and analyzed with an in-house Reynolds Average Navier-Stokes solver. The property of mesh alignment, mesh type and mesh distribution are analyzed based on the double cone configuration and the 69-degree Delta Wing Body configuration. The experimental data from wind tunnel are used as comparisons. The results show that the Mach cone aligned method has significant improvement to the accuracy for the prediction of sonic boom in nearfield. Both structured and un-structed mesh with a cylinder zone near the model can help to capture the distribution of the over-pressure, however, it is more convenient to use unstructured mesh when deal with complex configurations. The aspect ratio and circumferential spacing affect the captured shock wave peak value and the shapes. The accuracy of the prediction can be satisfied by carefully selecting these two factors. In the last, the summarized strategies for mesh generation are applied in the sonic boom prediction of a complex supersonic transport configuration, the Lockheed Martin 1021 to verify its practicability.

Near-Field Sonic-Boom Prediction and Analysis with Hybrid Grid Navier–Stokes Solver

A high-fidelity prediction of near-field sonic-boom signature is performed. Three benchmark cases (including an axisymmetric body, a simple delta wing–body, and a complete low-boom supersonic transport configuration) and the Lockheed Martin 1021 (LM1021) case are simulated with a hybrid grid Reynolds-averaged Navier–Stokes solver, HUNS3D, to predict near-field sonic-boom signature. Overall, the computational results agree well with the wind-tunnel experiment data. The influence of geometrical modification, grid size and distribution, spatial discretization schemes, and viscosity are studied and analyzed. The near-field signature of the LM1021 configuration propagating to the ground is analyzed by a linear wave approach.

Low-Boom/Low-Drag Design Optimization of Innovative Supersonic Transport Configuration

The generation of shock waves is inevitable in supersonic cruise, which results in the generation of wave drag as well as sonic boom on the ground. Some innovative concepts such as the supersonic biplane concept and the supersonic twin-body fuselage concept have been proposed recently to reduce the supersonic wave drag dramatically. In this study, the aerodynamic and sonic-boom performance of innovative supersonic transport wing–body configurations is discussed by numerical approaches. The wing section shape of a Busemann biplane wing/twin-body model is optimized at a design Mach number of 1.7 and angle of attack of 2 deg. The optimized results show the tradeoff between aerodynamic performance (lift/drag ratio) and sonic-boom performance (maximum overpressure of sonic-boom distribution). It is confirmed that aerodynamic performance is mainly affected by the thickness difference between the upper/lower wings. Then, the trailing-edge shape of the upper wing has also a significant influence on aerodynamic performance. On the other hand, sonic-boom performance is mainly affected by the lower surface shape of the lower wing, to make the pressure wave pattern of “expansion–compression”, which can attenuate the pressure waves propagating to the ground.

Aerodynamic/Sonic Boom Performance Evaluation of Innovative Supersonic Transport Configurations

In this research, biplane wing/twin-body fuselage innovative configurations are discussed as a next-generation supersonic transport candidate. Those aerodynamic performance as well as sonic boom performance are investigated by using numerical approaches. The aerodynamic performance is evaluated by inviscid compressible Euler equations using an unstructured mesh finite-volume method, and then the sonic boom performance is evaluated by an augmented Burgers equation. Thanks to the successful interactions of shock waves between the biplane wing as well as the twin-body fuselages, remarkable drag reduction and better sonic boom performance have been realized simultaneously at our design Mach number of 1.7. The superiority of the proposed supersonic transport configuration over conventional configurations is clearly demonstrated.

Fundamental Design Optimization of an Innovative Supersonic Transport Configuration and Its Design Knowledge Extraction

Fundamental Design Optimization of an Innovative Supersonic Transport Configuration and Its Design Knowledge Extraction

Biplane-Wing/Twin-Body-Fuselage Configuration for Innovative Supersonic Transport

In this paper, a biplane-wing/twin-body-fuselage configuration is discussed for advanced supersonic transport. The supersonic-biplane-wing concept has been proposed by Kusunose et al. (“Supersonic Biplane—A Review,” Progress in Aerospace Sciences, Vol. 47, No. 1, 2011, pp. 53–87), and a remarkable drag reduction in supersonic flow conditions has been demonstrated in the literatures. Motivated by the wave reduction effect of the supersonic biplane airfoils, a twin-body-fuselage concept has also been proposed by the present authors. Wave-drag reduction was achieved by an optimal twin-body-fuselage configuration. In this research, therefore, the biplane-wing concept and the twin-body-fuselage concept are merged to propose an innovative supersonic transport configuration. The wave-drag characteristics, as well as the wave-drag reduction, are illustrated by inviscid computational fluid dynamics computations at the design Mach number of 1.7. The inviscid (wave) drag of the proposed configuration is much smaller than that of the Sears–Haack single body with a tapered wing of a diamond–wedge airfoil. The increment of the skin-friction drag is also discussed for the proposed configurations. Based on the total drag analysis, the superiority of the biplane-wing/twin-body-fuselage configuration over the conventional configurations is clearly demonstrated.

Flow Simulation of an Supersonic Transport Configuration at Low-Speed and High-Lift Conditions

The flowfield around a supersonic transport configuration with high-lift devices at a low speed and high angle of attack was investigated by solving Reynolds-averaged Navier–Stokes equations. The configuration consisted of a fuselageandacrankedarrowwingwithleading-andtrailing-edge flaps.Numericalsimulationswereconductedand validated at conditions of the wind-tunnel test. Details of the flowfield at the design condition were analyzed using computational results. The effect of the high-lift devices on aerodynamic performance was discussed. The leadingedgevorticeswerereducedbothinsizeandinstrengthbydeflectingtheleading-edge flap,andthedragwasreduced. The trailing-edge flap increased the effective camber of the wing and improved the lift force. A typical leading-edge vortex flap was confirmed. It was shown that the aerodynamic performance could be improved by deflecting the leading- and trailing-edge flaps.

Investigations for Supersonic Transports at Transonic and Supersonic Conditions

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.

Vortex Behaviors of Rolled Supersonic Transport Configuration with Leading-Edge Flaps

Wind tunnel measurements were done to investigate the effect of leading-edge flaps on the rolling moment characteristics of the cranked arrow wing for the supersonic transport. Static rolling moment measurements, flow visualization studies, and crossflow velocity measurements by a particle image velocimetry were made at Re = 6.2 x 104. Static rolling moment measurements confirmed linear restoring moment at « < 16deg for the models both with and without flap deflection. When the outboard leading-edge flap is deflected 12 deg, rolling moment hysteresis is observed at the roll angle of 0 as 20 deg and α ≈ 20 deg. The experimental results indicate that the hysteresis is caused by different vortex breakdown behaviors on the inboard wing when the wing is rotated in the clockwise and counterclockwise directions.

Rolling Moment Characteristics of Supersonic Transport Configurations at High Incidence Angles

Wind-tunnel tests were performed to investigate rolling moment characteristics of a cranked arrow wing supersonic transport configuration at high incidence angles. Force measurements and surface static pressure measurements were conducted when the wing model was rolled statically at a Reynolds number of 9.21 x 10 3 based on the mean aerodynamic chord. To understand the behavior of the leading-edge separation vortices formed both on the inboard and outboard of the cranked arrow wing, a stereoscopic particle image velocimetry survey was performed at 55 and 83% root chord locations. As the wing roll angle was increased, linear stable rolling moments were observed at low incidence angles, whereas abrupt changes from stable to unstable rolling moments were observed at incidence angle of 20 deg. Variations in suction peaks of pressure distributions at the rear part of the wing contribute to destabilization of the rolling moment component. These variations in suction peaks that induce the unstable rolling moment are related to the vortex breakdown characteristics. Chordwise breakdown locations of the inboard and outboard vortices are different between the windward and leeward wings. This difference causes the sudden change of rolling moments.

Studies on Vortex Flaps with Rounded Leading Edges for Supersonic Transport Configuration

Wind-tunnel measurements were taken on a cranked arrow wing supersonic transport configuration with leading-edge vortex flaps. Force and surface pressure measurements were made at Reynolds number based on the wing mean aerodynamic chord from 9.2 × 105 to 3.8 × 10 6 . Two different flap cross sections (the originally designed nonrounded leading edge and the rounded leading edge) were tested. The purpose of the measurements is to clarify how the differences of the Reynolds number affect the flow around the rounded leading-edge vortex flaps and the flap performance. The wing with the rounded leading-edge vortex flaps indicated some benefit of the lift/drag ratio as compared with those of the nonrounded vortex flaps at a relatively high-lift coefficient greater than 0.3

超音速旅客機の離陸上昇性能について―前縁ボルテックス・フラップと後縁フラップの併用―

Improvements of the take-off and climb performance of the next generation supersonic transport (SST) are one of the key features for the SST development. Take-off and climb performances have been estimated for the cranked-arrow-wing SST configuration when the leading-edge vortex flaps and the trailing-edge flaps are deflected. Results show that the take-off distance and the balanced field length are reduced when the trailing-edge flaps are deflected, as expected. The thrust required for the constant climb gradient can be reduced when the leading-edge vortex flaps and the trailing-edge flaps are deflected at the same time.

Aerodynamic shape optimization of supersonic aircraft configurations via an adjoint formulation on distributed memory parallel computers

This work describes the application of a control theory-based aerodynamic shape optimization method to the problem of supersonic aircraft design. A high fidelity computational fluid dynamics (CFD) algorithm modelling the Euler equations is used to calculate the aerodynamic properties of complex three-dimensional aircraft configurations. The design process is greatly accelerated through the use of both control theory and parallel computing. Control theory is employed to derive the adjoint differential equations whose solution allows for the evaluation of design gradient information at a fraction of the computational cost required by previous design methods. The resulting problem is then implemented in parallel using a domain decomposition approach, an optimized communication schedule, and the Message Passing Interface (MPI) Standard for portability and efficiency. In our earlier studies, the serial implementation of this design method, was shown to be effective for the optimization of airfoils, wings, wing–bodies, and complex aircraft configurations using both the potential equation and the Euler equations. In this work, our concern will be to extend the methodologies such that the combined capabilities of these new technologies can be used routinely and efficiently in an industrial design environment. The aerodynamic optimization of a supersonic transport configuration is presented as a demonstration test case of the capability. A particular difficulty of this test case is posed by the close coupling of the propulsion/airframe integration.

Research on a sonic boom evaluation and low-boom and low-drag design

Kawasaki Heavy Industries, Ltd. has been researching on a sonic boom since 1990 toward the development of the next-generation SST (super sonic transport). A small chamber where a similar-sonic-boom could be created was made in 1993, and over 200 tests were carried out to see the reaction of people inside the chamber. As the result, it was found that A-weighted sound-pressure level (ASEL) and Stevens’ Mark VII perceived level (PL) were good metrics for boom evaluation, and the intolerable levels of ASEL and PL were studied. At the same time, an aerodynamic design technique was developed, which was produced by combining the low-boom theory of Darden (NASA) with the low-drag design method of super sonic aircraft. The validity of the design technique was verified by computational fluid dynamics (CFD) and wind tunnel testing. The effect of trim characteristics on a sonic boom has also been studied for more practical investigations. A boom propagation has also been calculated recently by using Hayes’ program for more understanding. The effects of weight, altitude, and Mach number on a propagating boom were investigated by parametric studies using the next-generation SST configurations in Japan.

Viscous flow past a nacelle in proximity to a flat plate

Introduction V ARIOUS conceptual designs for a commercial supersonic transport have been developed in recent years. Most of these configurations have been designed using linearized and modified linearized potential theory methodologies to define the vehicle shape and the aerodynamic characteristics. Other methods based on Whitham's modified linear theory^ have been used in the design and analysis of low-boom configurations. The combination of these methods has proven to be useful in the preliminary design stage of low-boom, high-speed civil transport configurations. In estimating the overall aerodynamic characteristics of a configuration with nacelles, the interference effects must be evaluated. The lift induced by nacelles on the wing lower surface adds to the equivalent body area distributions composed of volume and lift contributions. The equivalent body area rule defines Whitham's F-function, which in turn renders the far-field overpressure signatures of a supersonic aircraft. The computational technique described in Ref. 4 provides estimations of the loads imposed on a wing surface by nacelles mounted nearby. This numerical method, however, is impaired by its inability to account for certain nonlinear effects inherent in complicated flows. The qualitative analyses of wind-tunnel data for various aircraft models with nacelles have suggested that the interference pressures, calculated using the computer program of Ref. 4, might be underestimated in the extreme near field. Therefore, more accurate estimates of the extreme near-field conditions are required to improve the design of low-boom supersonic transport configurations. The purpose of the present study is to perform NavierStokes calculations in order to provide pressure distributions and force data on a flat plate with a nacelle in close proximity. The pressure distributions obtained using Navier-Stokes equations are then compared with the pressure distributions obtained using the linear theory of Ref. 4. Incorporating the nacelle-induced lift on the wing into the design process of a low-boom transport configuration will be included in a followup study. The Navier-Stokes equations are solved by an implicit, approximately factored, finite volume, upwind algorithm. ) Baysal et al. have used this algorithm to compute a supersonic flow past an ogive-nose-cylinder at zero-deg angle of attack near a flat plate on a composite mesh of over-

Three-dimensional time-marching aeroelastic analyses using an unstructured-grid Euler method

Modifications to a three-dimensional, implicit, upwind, unstructured-grid Euler code for aeroelastic analysis of complete aircraft configurations are described. The modifications involve the addition of the structural equations of motion for their simultaneous time integration with the governing flow equations. The paper presents a detailed description of the time-marching aeroelastic procedures and presents comparisons with experimental data to provide an assessment of the capability. Flutter results are shown for an isolated 45-deg swept-back wing and a supersonic transport configuration with a fuselage, clipped delta wing, and two identical rearward-mounted engine nacelles. Comparisons are made between computed and experimental flutter characteristics to assess the accuracy of the aeroelastic results.

Minimum Mass Sizing of a Large Low-Aspect Ratio Airframe for Flutter-Free Performance

A procedure for sizing an airframe for flutter-free performance is demonstrated on a large, flexible supersonic transport aircraft. The procedure is based on using a two-level reduced basis or modal technique for reducing the computational cost of performing the repetitive flutter analyses. The supersonic transport aircraft exhibits complex dynamic behavior, has a well-known flutter problem and requires a large finite-element model to predict the vibratory and flutter response. Flutter-free designs are produced with small mass increases relative to the wing structural weight and aircraft payload. In view of the ability of the resizing procedure to handle this supersonic transport configuration, it seems likely that the method could be used for many other aircraft.

Evaluation of Techniques for Predicting Static Aeroelastic Effects on Flexible Aircraft

An experimental evaluation of analytical techniques for predicting the longitudinal stability characteristics of a large flexible aircraft is presented. Analytical methods based on both the modal approach and stiffness influence coefficients are used to predict the aerodynamic characteristics of a flexible airplane. These methods are then applied to a flexibly scaled model of a supersonic transport configuration. Comparisons between wind-tunnel data, the modal approach, and calculations based on stiffness influence coefficients are presented over the Mach number range from M = 0.6-2.7. The results of this study show good agreement for most cases between analyses and experiment. Nomenclature a.c. = aerodynamic center position Aa.c. = shift in aerodynamic center position due to flexibility, positive forward b — wing span c = local chord measured streamwise c = reference chord CL = lift coefficient, lift/gS CLx = lift-curve slope, 8CL/dot at a = 0° CL | a =o = lift coefficient at a = 0° CLq = lift coefficient due to pitching, 8CL/d(9c/2V) CL5e = elevator effectiveness in lift, dCL/88e