Fighter Configuration Research Articles

Performance augmentation of a 60-degree delta aircraft configurationby spanwise blowing

Spanwise blowing (SWB) over the wing and canard of a close-coupled-canard, 60-deg delta fighter-aircraft configuration was investigated experimentally in low-speed flow at angles of attack up to 60 deg and yaw angles of up to 36 deg. Significant improvement in lift-curve slope, maximum lift, drag polar, and lateral/directional stability was found, enlarging the usable flight envelope beyond its previous low-speed/maximum-lift limit. It was shown that SWB can achieve the same lift augmentation produced by a canard, without the drag penalty. Contrary to previous experience with 60-deg swept wings, the efficiency of the lift augmentation by SWB was relatively high and was found to increase with increasing jet-momentum coefficient on the close-coupled-canard configuration. Interesting and promising possibilities of obtaining much higher efficiencies with swirling or multiple nonaligned jets were indicated.

Euler equations analysis of the initial roll-up of aircraft wakes

A new approach has been developed to study the initial roll-up of aircraft wakes. This approach is based on the solution of the three-dimensional Euler equations subject to inflow and initial conditions predicted by a vortex-lattice method. This procedure allows for the solution of this complex problem without any empirical inputs. However, valid solutions are limited to those flowfields generated by wings which are free of flow separation. Results are presented for four configurations: a rectangular wing with part span flap, a rectangular wing, a typical transport-type configuration and a fighter configuration. In addition, an assessment of an 'unsteady 2-D analogy' was undertaken using the 2-D time accurate Euler equation solutions in crossflow planes. In general, predictions from the 2-D unsteady analogy coincided with the 3-D results which were found to be in good agreement with available experimental data.

Unsteady aerodynamic modeling of a fighter wing in transonic flow

A numerical method for predicting steady and unsteady aerodynamic flows about aircraft wing configurations is presented. The numerical procedure solves the three-dimensional full-potential equation by a strongly implicit approximate factorization algorithm. Steady-flow analyses are obtained by relaxation, while unsteady analyses are calculated by time-accurate marching. In the present work, calculated steady, quasisteady, and unsteady wing surface pressures are presented for an F-5 fighter configuration and compared to experimental data for subsonic and transonic flight conditions.

A Foward-Swept-Wing Fighter Configuration Designed by a Transonic Computational Method

A Foward-Swept-Wing Fighter Configuration Designed by a Transonic Computational Method

Application of NLR's calculation methods to transonic flow about oscillating wings

Unsteady and quasisteady aerodynamic loads in the transonic flow domain have been obtained from the numerical simulation methods of the National Aerospace Laboratory (NLR) for transonic potential flow. The unsteady results have been obtained with the time-linearized finite volume method FTRAN3, embedding a socalled field panel method which accounts for proper radiation of signals toward infinity. The mean steady flowfields and the quasisteady results have been obtained from the application of XFLO22—the NLR system for the simulation of steady transonic flow about wings and wing bodies. The calculations have been performed for several wing planforms and the results are compared with experimental data. A correlation is made between the quasisteady results of XFLO22 and FTRAN3. Flutter applications to a transonic transport wing model and a fighter configuration have been made.

Computation of supersonic flows over three-dimensional configurations

An aerodynamic prediction technique based on the steady form of the full-potential equation has been applied to a variety of three-dimensional supersonic flow problems exhibiting embedded subsonic regions. A conservative switching scheme is employed to transition from the supersonic relaxation algorithm, and vice versa. Numerical solutions are obtained for a number of complex configurations, including advanced tactical fighter, Langley canard-wing fighter configuration, isolated shuttle orbiter, and mated shuttle orbiter configuration with external tank. The computed results are in good agreement with available experimental data.

Impact of fuselage incidence on the supersonic aerodynamics of two fighter configurations

The results of experimental and theoretical investigations into the effect of fuselage upwash on fighter aircraft wing performance are reported. Wind tunnel trials were performed on 4 percent scale models of two supersonic fighters. The trials were run at Mach 1.6-2.0, an Re of 2,000,000 and at angles of attack (AOA) of -4 to 20 deg. Measurements were made of lift, drag and pitching moments. Two linearized theory supersonic aerodynamic prediction codes, PAN AIR and the SDAS lift analysis, were used to predict the same aerodynamic coefficients. The fuselage AOA augmented the lift and pitching moment at 0, 2 and 5 deg. The contribution mainly arose from the fuselage-induced upwash. The PAN AIR code gave superior data for the fuselage aerodynamics and effects, although both codes accurately predicted the overall lift and moment increments due to the fuselage AOA.

Approach and landing technologies for STOL fighter configurations

Providing next-generation fighter aircraft with short takeoff and landing capability will require effective highlift systems, methods of providing adequate longitudinal trim and control, and thrust reversers compatible with advanced configurations. Recent wind tunnel tests in the NASA Langley 4- x 7-m tunnel have provided the opportunity to investigate some promising technologies for providing this short takeoff and landing capability. These tests include investigations of the high-lift contribution of spanwise blowing on a trailing-edge flap system, the trim capability of a blown high-lift canard, and the deceleration of a thrust reverser for ground roll reduction. This paper reviews the results of each investigation and examines the integration of these technologies on a common advanced fighter configuration to determine their compatibility and the resulting overall performance of the configuration.

Flowfield investigation of a supercruise fighter model

A NASA Langley investigation was conducted in the 16-foot Transonic Tunnel to survey the flow field around a model of a Supersonic cruise fighter configuration. In this investigation, a model of a supersonic cruise fighter configuration formerly utilized in afterbody-nozzle performance investigations was surveyed with a single, multiholed probe to determine local values of angle of attack, side flow, and Mach number. The investigation was conducted at Mach numbers of 0.6, 0.9, and 1.2 at angles of attack from 0 to 10 deg. The purpose of the investigation was to provide a data base of experimental data for use in verification of theoretical methods, and to compare the experimental data with predictions from currently available theoretical techniques. Results from this investigation show that local angles of attack were generally greater than free stream above the wing and generally less than free stream below the wing. Also there were large spanwise gradients above the wing at the higher angles of attack. The comparisons of experimental data with theoretical predictions show that the theoretical techniques give a qualitative estimate of the flow-field but will require much work to give good quantitative results.

Treatment of supersonic flows with embedded subsonic regions

A nonlinear method based on the full potential equation in conservation form, cast in an arbitrary coordinate system, has been developed to treat predominantly supersonic flows with embedded subsonic regions. This type of flow field occurs frequently near the fuselage/canopy junction area and wing leading-edge regions for a moderately swept fighter configuration. The method uses the theory of characteristics to accurately monitor the type-dependent flowfield. A conservative switching scheme is developed to handle the transition from the supersonic marching algorithm to a subsonic relaxation procedure, and vice versa. An implicit approximate factorization scheme is employed to solve the finite differenced equation. Results are shown for a few configurations, including a wing/body/wake realistic fighter model having embedded subsonic regions.

Free-flight and wind-tunnel data for a generic fighter configuration

A comparison of free-flight spark range and wind-tunnel data for a generic fighter configuration (standard dynamics model) is presented. The aerodynamic tests were conducted from 0.3 to 1.3 with the primary comparisons at the transonic Mach numbers. This paper shows the comparison of the zero angle-of-attack coefficients and derivatives since only relatively small oscillation amplitudes were achieved during the free-flight tests. However, it is believed that this comparison represents good agreement between these two sets of data obtained in different facilities using different test mechanisms and techniques.

Trimming advanced fighters for STOL approaches

Recent low-speed wind-tunnel tests of two powered advanced fighter configurations with short takeoff and landing (STOL) capability provided the opportunity to examine two different methods of generating the nose-up pitching moments required to maintain longitudinal trim during approach conditions. The pitching moments were generated by a small direct-lift nose jet on a supersonic cruise configuration and by a blown high-lift canard on an advanced multirole fighter configuration. Results of these investigations indicate that at nominal approach conditions (i.e., military power, main nozzle deflections about 40 deg and angle of attack about 15-16 deg) the nose jet and the blown high-lift canard provide sufficient nose-up pitching moment to allow the respective configurations to be trimmed. Neither configuration, could be trimmed at approach conditions without these devices.

Piloted simulation of hover and transition of a vertical attitude takeoff and landing aircraft

Piloted simulation studies of candidate control systems for VATOL aircraft were conducted on a six degree of freedom simulator. Hover and transitions from wing-born to hovering flight were performed, with and without turbulence, on a representative high performance fighter configuration. Deflection of the rear engine nozzle provided pitch and yaw control moments in concert with reaction controls for roll. Unique motion cues in hover result from the vertical displacement of the cockpit and the thrust vectoring nozzles. Abundant control power available with moderate engine nozzle deflection combined with rate feedback for stability augmentation provided very satisfactory control.

Departure susceptibility and uncoordinated roll-reversal boundaries for fighter configurations

A fighter's response to a severe application of controls is predicted as a function of its static lateral and directional stability derivatives and lateral control characteristics. This was accomplished by developing two boundaries which identify the onset of roll reversal and departure in terms of the lateral-directional characteristics. These boundaries demark four airplane response regions, i.e., commanded maneuver, commanded maneuver accompanied by an alpha excursion, uncoordinated roll reversal, and departure susceptibility. The responses within the latter two regions become progressively more severe the further the lateral-directional characteristics are displaced from the onset boundary values. The boundaries were developed and their sensitivity to other variables was determined through large-angle, six-degree-of-freedom parametric computer studies. Since they were found to be applicable for most fighter mass distributions and for a large range of pitch characteristics, it was concluded that they could be used for specification, design, and evaluation purposes.

Analysis of Thrust-Induced Effects on the Longitudinal Aerodynamics of STOL Fighter Configurations

An analysis of the thrust-induced longitudinal aerodynamic characteristics of three fighter-type configurations is presented. A brief discussion of the take-off and landing requirements for the next generation fighter aircraft leads to the conclusion that advances in lift coefficient and thrust reversing will be required to allow short-field operation. Typical power-on longitudinal aerodynamic data for the three fighter configurations, indicating different approaches to meeting the high-lift coefficient requirement, are discussed. Thrust reversing is not addressed in this paper. The power-on data are analyzed to determine what power effects are present, that is, direct thrust, boundary-layer control, vortex flows, or supercirculation. The results of the analysis indicate, for the configurations studied, that induced effects are small compared to direct thrust and that boundary-layer control, rather than leading-edge vortex flows or supercirculation, is the only significant thrust-induced effect.

Aerodynamic Characteristics of Fighter Configurations During Spin Entries and Developed Spins

The NASA Langley Research Center is currently conducting a stall/spin research program to define fighter aerodynamics applicable during spin entries and developed spins and to develop analytical methods to use such measured aerodynamics for theoretically calculating these spin motions. Some static, forced-oscillation, and continuous rotation aerodynamic data have been measured for several current fighter models over a large poststall angle-of-attack range. This paper discusses these aerodynamic data and illustrates both the extremely nonlinear dependence of such data on several variables and the correlation that exists between the three types of measured aerodynamics. The current analytical methods for using these areodynamics to calculate spin entry and developed spin motions are discussed and correlated with experimentally obtained spins.

Deflected Thrust Effects on a Close-Coupled Canard Configuration

The effects of deflected thrust on the stability and performance of a close-coupled canard fighter configuration are presented. These results were obtained at low speeds in the Langley V/STOL tunnel. Transonic as well as low-speed results are also presented for an unpowered close-coupled canard and a supercruiser configuration. The V/STOL tunnel data indicate an increase in maximum lift and reductions in drag due to lift with the addition of two-dimension al vectored thrust at the wing inboard trailing edge. The longitudinal pitchup associated with the unpowered configuration at higher angles of attack was significantly reduced with power.

Advanced Fighter Controls Flight Simulator for All-Systems Compatibility Testing

McDonnell Aircraft Company has developed a flight simulator to serve as a test bed for control systems for advanced fighters. The cockpit of the simulator is adaptable to any fighter cockpit configuration and is located on the nose of a hydraulic, electrical and control system test bed. This paper contains a description of how this unique simulation facility was set up, checked out and operated in the all-systems compatibility testing of the F-4 fly-by-wire Survivable Flight Control System (SFCS). Actual SFCS flight hardware and electronics were mated to an existing control system mock up and to a digital computer simulating F-4 flight dynamics. The simulator flights included takeoff, landing, air-to-air and air-to-ground target tracking and system failures, as well as the usual flight test handling quality maneuvers. In addition, normal and emergency procedures were developed. Thus, through testing and training with the simulator, control system and interface problems were corrected and a high level of confidence in the design of the SFCS was established before an SFCS equipped aircraft was flown. These accomplishments, and a capability for rapid response in testing problems and modifications developed during flight test were major factors in keeping the program on schedule and on cost.

Transonic Navier-Stokes flow computations over wing-fuselage geometries

The steady-state transonic viscous flows over two wing-fuselage configurations are solved numerically. The 3D compressible thin-layer Navier-Stokes equations are solved by means of multistage timestepping similar to the Runge-Kutta method. The timestepping technique is applied with a multigrid scheme to make the transonic Navier-Stokes computations using the Baldwin-Lomax turbulence for closure and converging the solution on a grid of 1.52 million grid points with a C-O topology. Both generic and fighter wing-fuselage geometries are solved, and the results are compared to experimental studies. Solution convergence is found to be fast, and the fine grid generates accurate results for both configurations by means of the Baldwin-Lomax turbulence model. Results from the Johnson-King model calculations produce results that are more consistent with experimental results for the fighter configuration.

Reingestion characteristics and inlet flow distortion of V/STOL lift-engine fighter configurations.

Exhaust gas reingestion and inlet flow distortion in V/STOL lift engines, discussing static and wind tunnel testing using fighter configuration models