Aircraft Forebody Research Articles

Hypersonic pressure-controllable bump based on an improved permeable-boundary method

This paper presents a novel design method of pressure-controllable bump for hypersonic aircraft forebody. The new developed method can effectively resolve the tradeoff among boundary layer diversion, flow uniformity, and external drag reduction. The classical permeable-boundary method is improved by coupling with the radius-based function, through which the prescribed surface pressure distribution can inversely generate the bump. The improved permeable-boundary method is available for 3D unstructured mesh which increases the solution efficiency and robustness. The accuracy of the method is evaluated through the comparison with experimental results. In addition, five principles to arrange the pressure distribution are proposed. Then, a new pressure-controllable bump is designed. Compared with the typical streamline-tracing bump, the new bump is 44.9% lower in height while diverting identical low kinetic energy flow. According to the space occupied with high kinetic energy flow, the uniform region to preset the inlet of the new bump is 58.0% wider than that of the typical bump. Moreover, the usage efficiency of high kinetic energy flow is 1.5% higher. The new bump is less convex and its equivalent lift-to-drag ratio is two times greater than the typical bump. This research confirms that the new pressure-controllable bump shows better integrating capacity with the hypersonic inlets than the typical bump. The current design method is based on an inviscid condition; the method considering the viscous effects will be further studied in the future.

A pressure-controllable bump based on the pressure-ridge concept

An innovative type of pressure distribution for the hypersonic aircraft forebody (bump) is presented, and this design is based on the newly established known as pressure ridge (PR) flow mechanism. Studies on the low-kinetic-energy fluid over the aircraft surface are summarized. Then, the challenges of the inlet–airframe integration at high speeds are discussed. The flow structure around the bump is analyzed in detail to eliminate the side-embedded shock (SES) effect at the inlet entrance. The concept of PR is proposed to improve the overall aerodynamic characteristics of the bump, namely, the boundary layer removal, the reduction of external drag, and the streamline direction at bump end-section. On the basis of the developed inverse method to generate the bump surface by the prescribed pressure distribution, the improved PR-derived bump is designed and numerically compared with the typical pressure-controllable bump (PCB) while the identical leading edge profile is imposed. Results demonstrate that the PR creates outward and inward pressure gradients. The removed amount of boundary layer increases with the increase in outward pressure gradient. Meanwhile, the inward pressure gradient determines the streamline pattern after the bump. The PR-derived bump is 29.2% lower in height than the typical one by using a proper outward pressure gradient. The uniform area of the new bump is 30% wider than that of the typical PCB. The near-wall streamlines of the new bump are adjusted from expanding to the parallel, thereby relieving the side-compression. Changing the location, width, and peak value of the PR can lead to great flexibility in the design and optimization of aircraft forebody subjected to hypersonic flow.

The Stressed State of a Stiffened Conical Shell with Thermal Protective Coating with Temperature-Dependent Properties

The equations of the mathematical model are solved in terms of special functions. The results for the design scheme of the aircraft forebody are obtained with a guaranteed accuracy by the stable method of functional normalization.

Design method of internal waverider inlet under non-uniform upstream for inlet/forebody integration

A novel Bump-integrated three-dimensional internal waverider inlet (IWI) design method is presented for high-speed inlet/forebody integration. The low-kinetic-energy (boundary layer) flow generated by a blunted leading-edge and forebody boundary layer represents an extreme challenge in the integration of aircraft forebody and inlet. In this method, such an inlet's flowpath is divided into the entrance shockwave segment, the isentropic compression segment and the isolator. First, a three-dimensional inverse method of characteristics (3D-IMOC) is developed to obtain a compression surface that can generate a requested entrance shock wave in non-uniform upstream flow. This configuration realizes the integration of IWI and aircraft fuselage by incorporating a Bump to remove most of the boundary layer flow. This is followed by a three-dimensional, isentropic compression flow-path with cross sectional areas conforming to the specified Mach number distribution. Finally, a new three-dimensional Bump-integrated IWI was tested in M = 6 wind tunnel, under a rather thick boundary layer upstream flow (37% height of inlet entrance). Both of the experimental data and numerical simulation results show that, the new method of IWI and Bump can overcome serious boundary layer flow problems and improve the inlet performance.

The engineering model of non-equilibrium flow

A physico-mathematical model of non-equilibrium, primarily, hypersonic flow is proposed. The model is meant for calculating the flow around the aircraft forebody. By taking as an example the solution of the problem on the flat shock wave profile, it is shown that the model being proposed allows the traditional Navier-Stokes-Fourier model to be improved.

A neural network approach to the calibration of a flush air data system

AbstractA flush air data system uses surface pressure measurements to obtain speed and aerodynamic orientation of a flight vehicle. This paper investigates the use of a neural network to calibrate the air data system on a low-observable aircraft forebody with 22 pressure tappings. Wind tunnel data were obtained for between 0 and 25° angle of attack, 0 to 10° sideslip for speeds of 32, 37 and 45ms-1. Experimental data were used to train multilayer perceptron neural networks. A calibration accuracy of 0·322ms-1, 0·811° and 0·552° for speed, alpha and beta respectively was achieved just using the first five tappings on the nose cone. Increasing the number of tappings used as inputs to the neural network reduces the calibration error. A neural network with 22 inputs gives a best accuracy of 0·095ms-1, 0·15° and 0·085° for speed, α and β respectively. Trained networks show poor robustness to single sensor failures. Robustness is improved by preprocessing input data with an autoassociative network or by introducing sensor redundancy.

Review of forebody-induced wing rock

Wing rock at moderate angles of attack is often generated by dynamic wing stall; however, the wing rock of most concern is generated by the vortices shed from the aircraft forebody at high angles of attack. The capability to understand and predict forebody-induced wing rock is a prerequisite for the successful design of agile aircraft and missiles that are required to perform maneuvers at high angles of attack. Available information on the wing rock experienced by advanced aircraft is reviewed and an explanation of the flow physics causing the wing rock is presented.

Processing flight test data

The calibration of measurements of the four critical parameters that define an aircraft's air data state, the state of an aircraft relative to the air mass through which it is flying, is discussed. The parameters are altitude, airspeed, angle-of-attack, and angle-of-sideslip. The measurements are taken by external probes, usually located on the aircraft's forebody, and are affected by the aircraft contours, which tend to disturb the local airflow across the probes. It is shown that using an expert system to calibrate the data improves the process. >

Numerical solutions of the Euler equations governing axisymmetric and three-dimensional transonic flow

SummaryA computational method is presented for solving numerically the three dimensional Euler equations for transonic flow around practical aircraft forebodies. The Euler solver is pseudo-time dependent, split and cast in finite volume form. Shock waves are captured crisply without the need for additional smoothing by means of an operator-switching facility which more accurately reflects the direction of propagation of signals.The method is illustrated by examples of computed external, axisymmetric flows and some, simulated, realistic aircraft fore-bodies. The computational meshes employed in the three dimensional cases are essentially of cylindrical polar, flow conforming, type and relatively coarse. Closer attention to the mesh generation is expected to refine the results presented here. The method is versatile, robust, and holds promise for treating complex three dimensional geometries within economically viable run times.

Control of Forebody Vortex Orientation to Enhance Departure Recovery of Fighter Aircraft

A combined experimental and analytical study was undertaken to develop active blowing concepts to control the asymmetric orientation of the vortex system emanating from an aircraft forebody at high angles of attack. The objective of the study was to utilize the side force associated with asymmetric vortices, in a controlled manner, to enhance the capability of a fighter to recover from a departure from controlled flight. Results from water tunnel and wind tunnel tests show that a small amount of tangential blowing along the forebody can effectively alter the forebody vortex system and generate large restoring yawing moments. Six degree-of-freedom digital simulation results show this concept can substantially enhance departure recovery characteristics of fighter aircraft having long, slender forebodies.

Water Tunnel Flow Visualization: Insight into Complex Three-Dimensional Flowfields

Water tunnel facilities and flow visualization techniques have been developed at Northrop to provide high quality visualization of vortex interactions at high angles of attack. Results have provided considerable insight into highly complex three-dimensional flowfields generated by contemporary fighter aircraft. Studies have been made of the leading edge vortex systems generated by wing leading edge extensions (LEXs) typical of current subsonic-transonic fighter aircraft, high angle of attack aerodynamic asymmetries associated with the vortex system developed on long slender forebodies, and forebody and wing/LEX vortex interactions characteristic of highly maneuverable aircraft with hybrid wing planforms and slender forebodies. Qualitative results are in excellent agreement with existing subsonic wind tunnel data. The water-to-air analogy has been verified, that is, aircraft forebody and wing/LEX vortex systems, vortex system interactions, and the downstream influence on flow characteristics exhibited in air at high Reynolds numbers can be simulated at low Reynolds numbers in the water tunnel.

Inlet Boundary-Layer Shapes on Four Aircraft Forebodies at Mach 6

Boundary-layer shape measurements at the engine inlet on four different hypersonic aircraft forebody designs (with no engine on the forebody) are reported. The measurements provide a qualitative assessment of the effectiveness of various forebody geometries as engine inlet precompression surfaces. The designs, tested in a hypersonic tunnel at Mach 6 and a nominal freestream Reynolds number of 30,500,000, included a semiconical forebody, a configuration similar to a slab delta wing, a conical nose blended into a flat surface, and a conical, complex forebody shape. Boundary layer height as a function of forebody compression is shown for each design.

Computational investigation of a pneumatic forebody flow control concept

The effectiveness of a tangential slot blowing concept for generating lateral control forces on an aircraft forebody is analyzed using computational fluid dynamics. The flow about a fighter forebody is computed using a multiple-zone, thin-layer Navier-Stokes code. Tangential slot blowing is modeled by the use of an actuator plane. The effects of slot location and slot length on the efficiency of the system are analyzed. Results of the study indicate that placement of the slot near the nose of the aircraft greatly enhances the efficiency of the system, while the length and circumferential location of the slot are of secondary importance. Efficiency is defined by the amount of side force or yawing moment obtained per unit blowing coefficient. The effect of sideslip on the system is also analyzed. The system is able to generate incremental changes in forces and moments in flows with sideslip angles up to 10 deg comparable to those obtained at zero sideslip. These results are used to determine a baseline configuration for an experimental study of the tangential slot blowing concept.