Aircraft Geometry Research Articles

Prismatic grid generation for three-dimensional complex geometries

The paper describes generation of semiunstructu red prismatic grids for three-dimensional complex geometries. A new method of generation is presented, which couples both algebraic and elliptic steps. This combination yields low computation cost, as well as smoothness of the grid. Main advantages of the developed grid generator are its direct control of grid orthogonality and spacing, as well as its generality for treatment of three-dimensional geometries. The method marches a surface consisting of triangular faces away from the body surface. Employment of prismatic elements is a relatively new approach toward three-dimensional, complex geometry numerical simulations of viscous flows. A body surface is covered with triangles, which gives the geometric flexibility, whereas the structure of the mesh in the direction normal to the surface provides better resolution of the viscous regions. An F-16A aircraft geometry was included in the applications to investigate efficiency and to demonstrate robustness of the method in handling relatively complex topologies. Grid generation time for the entire aircraft domain was of the order of minutes on a workstation.

Calculation of particle concentration around aircraft-like geometries

A computationally-efficient method is presented to calculate local particle concentration enhancements resulting from potential fluid flow around an idealized aircraft fuselage and wing. The geometries chosen for study are a 10:1 prolate ellipsoid at 0° angle of attack and a Joukowski airfoil at 0° and 5° angles of attack, for which potential flow analytic solutions are known. The collection efficiency of and surface concentration on a cylinder in potential flow are also considered for algorithm verification. Particle concentration is calculated along particle pathlines by a mixed Eulerian-Lagrangian technique developed by Fernandez de la Mora and Rosner (1981, Fernandez de la Mora, J. F. and Rosner, D. E., Physico Chem. Hydro. 2, 1). Ordinary differential equations for particle position, velocity, and concentration are integrated numerically by a variable order, backward difference algorithm. The calculations show the creation of regions of increased concentration near objects, and of particle-free shadow zones downstream. The magnitudes of the concentration disturbances are greatest at intermediate Stokes numbers (0.1–1.0) where inertia and drag are equally dominant. Samplers placed in these regions of enhanced particle concentration may not provide accurate concentration measurements. Ultimately, this approach could be included with detailed flow solutions about specific aircraft geometries to provide guidance in locating samplers in regions of acceptably small concentration deviations.

Regional fanjet aircraft optimization studies

This article illustrates the use of optimization methods in the development of a new regional aircraft. These aircraft are more sensitive to changes in operational requirements than other types due, in part, to their high zero-fuel weight ratio. For such aircraft it is essential that the information possible be available to the designers in the early project definition phases. The identification of optimum aircraft configurations and mission characteristics for various operational and configurational options constitutes a vital part of this knowledge. A series of industrially- related design studies are presented. These include 1) the selection of the baseline config- uration; 2) parameter sensitivity investigations; 3) analysis of aircraft and engine stretch potential; and 4) generalized (rubber engine) designs. This article concludes with a discussion on the merits of optimization studies in aircraft project design and offers some suggestions for changes to the strategies adopted. I NITIAL aircraft project design always presents difficult decision areas to designers. Once a particular aircraft con- figuration has been finalized it is too expensive to introduce major changes. Considerable pressure is exerted on the proj- ect team to get it right. The situation is further complicated by the changing technical and commercial environment sur- rounding the design. Some of the possibilities for change in- clude the choice of engine supplier, improvements in engine performance, introduction of new airframe and aircraft sys- tems, technology improvements, the emergence of new com- petitor aircraft, and changing operational and regulatory frameworks. The designers are therefore concerned with the specification of the initial design and the strategy for future aircraft development. The chosen aircraft configuration is un- likely to be exactly matched to the initial operational speci- fication but it must not be compromised to an extent that the first design i§ uncompetitive. It is within this context that the value of the optimum design method should be judged. The industrial design team places strong emphasis on com- ponent design aspects. These are pursued in fine detail and constitute substantial technical effort. Past experience has shown that careful control of the aircraft detail design in the early stages offers the method of avoiding technical difficulties later in the design cycle. Broader-based optimization studies can be considered as complementary to this detail design ef- fort. Together they provide both micro- and macro-design scenarios which help the designers in their difficult choice of initial aircraft specification and in subsequent developments. In the early stages of the design when the aircraft geometry and performance criterias have not been fixed, the optimi- zation studies are used to identify the absolute best con- figuration with respect to various operational specifications. This enables the designers to judge the sensitivity of the design with respect to variations of the aircraft geometrical and per- formance parameters. Once the baseline design has been es- tablished the optimization studies are used to evaluate the tolerance available in the selection of aircraft operational pa- rameters. Configurational features (e.g., rear-engine posi- tion) can also be assessed to show the quantifiable penalties/

The modelling of aerodynamic flows by solution of the euler equations on mixed polyhedral grids

AbstractIn this paper, an algorithm for obtaining solutions to the compressible Euler equations on mixed polyhedral grids is described. These grids, which are, in general, composed of hexahedral, pentahedral and tetrahedral elements, are used in the modelling of aerodynamic flows over complex three‐dimensional geometries. The hexahedra are grouped into regular blocks, while the other elements are part of irregular grid regions, giving rise to the term “hybrid grid approach”. The interconnection of the grid elements is defined by an efficient data structure, which supplies the required information to drive the Euler algorithm. In fact, cell‐vertex and cell‐centre spatial discretization variants of this algorithm are detailed, along with an explicit time‐marching scheme for obtaining steady‐state solutions. Results on two geometrically simple configurations provide an evaluation of these solution techniques, whilst the focus on a third, complex configuration, demonstrates the potential of the method in addressing the flow over general aircraft geometries.

Calculation of steady and unsteady pressures on wings at supersonic speeds with a transonic small-disturbance code

A transonic unsteady aerodynamic and aeroelasticity code has been developed for application to realistic aircraft configurations. The new code is called CAP-TSD which is an acronym for Computational Aeroelasticity Program - Transonic Small Disturbance. The CAP-TSD code uses a time-accurate approximate factorization algorithm for solution of the unsteady transonic small-disturbance equation that is efficient for solution of steady and unsteady transonic flow problems including supersonic freestream flows. The new code can treat complete aircraft geometries with multiple lifting surfaces and bodies. Applications to wings in supersonic freestream flow are presented. Comparisons with selected exact solutions from linear theory are presented showing generally favorable results. Calculations for both steady and oscillatory cases for the F-5 and RAE tailplane models are compared with experimental data and also show good overall agreement. Selected steady calculations are further compared with a steady flow Euler code.

Water droplet impingement on airfoils and aircraft engine inlets foricing analysis

This paper includes the results of a significant research program for verification of computer trajectory codes used in aircraft icing analysis. Experimental water droplet impingement data have been obtained in the NASA Lewis Research Center Icing Research Tunnel for a wide range of aircraft geometries and test conditions. The body whose impingement characteristics are required is covered at strategic locations by thin strips of moisture absorbing (blotter) paper and then exposed to an airstream containing a dyed-water spray cloud. Water droplet impingement data are extracted from the dyed blotter strips by measuring the optical reflectance of the dye deposit on the strips with an automated reflectometer. Impingement characteristics for all test geometries have also been calculated using two recently developed trajectory computer codes. Good agreement is obtained with experimental data. The experimental and analytical data show that maximum impingement efficiency and impingement limits increase with mean volumetric diameter for all geometries tested. For all inlet geometries tested, as the inlet mass flow is reduced, the maximum impingement efficiency is reduced and the location of the maximum impingement shifts toward the inlet inner cowl.

Application of Computer Graphics to Performance Studies of Missile Warheads

Intercept geometry of target aircraft and missiles play an important role in determining the effectiveness of the warhead. Factors such as fragment spatial distribution profile, damage capabilities, target and missile characteristics have been considered and visualised through computer graphics and optimum intercept intercept angles have been arrived. Computer graphics has proved to be an important tool to enhance perception and conceptual design capabilities in the design environment.

Further analysis of wing rock generated by forebody vortices

Further analysis of wing rock generated by forebody vortices has revealed a richness of flow phenomena that can produce the rather violent forms of wing-body rock observed experimentally. Relatively mild changes of the aircraft geometry cause different flow mechanisms to become the dominant ones, dictating the nature of the wing rock. The general observation is that the wing rock generated by vortices from a slender forebody not only is much more severe than slender wing rock, generated by asymmetric, leading-edge vortices, but it is also the type of wing rock to which most current and future advanced aircraft will be exposed.

Numerical simulation of the Navier-Stokes equations for an F-16A configuration

The results of a numerical simulation are presented for the steady flow over an F-16A aircraft configuration at a freestream Mach number of 1.2, a Reynolds number of 12.75 x 10, and an angle of attack of 6 deg. The three-dimensional Navier-Stokes equations in mass-averaged variables were integrated numerically using the MacCormack explicit algorithm with an algebraic turbulence model to provide closure of the system of equations. The grid structure, boundary conditions, and solution procedure are discussed in detail for this complex aircraft geometry. The results are then compared to experimental data in terms of surface pressure coefficients, lift coefficient, and drag coefficient with reasonable agreement. Finally, details of the flow are discussed, such as the strake vortex and the wing vortex structures.

Unsteady transonic flow calculations for realistic aircraft configurations

A transonic unsteady aerodynamic and aeroelasticity code has been developed for application to realistic aircraft configurations. The new code is called CAP-TSD which is an acronym for Computational Aeroelasticity Program - Transonic Small Disturbance. The CAP-TSD code uses a time-accurate approximate factorization (AF) algorithm for solution of the unsteady transonic small-disturbance equation. The AF algorithm is very efficient for solution of steady and unsteady transonic flow problems. It can provide accurate solutions in only several hundred time steps yielding a significant computational cost savings when compared to alternative methods. The new code can treat complete aircraft geometries with multiple lifting surfaces and bodies including canard, wing, tail, control surfaces, launchers, pylons, fuselage, stores, and nacelles. Applications are presented for a series of five configurations of increasing complexity to demonstrate the wide range of geometrical applicability of CAP-TSD. These results are in good agreement with available experimental steady and unsteady pressure data. Calculations for the General Dynamics one-ninth scale F-16C aircraft model are presented to demonstrate application to a realistic configuration. Unsteady results for the entire F-16C aircraft undergoing a rigid pitching motion illustrated the capability required to perform transonic unsteady aerodynamic and aeroelastic analyses for such configurations.

Grid generation and inviscid flow computation about a cranked-wingedairplane geometry

An algebraic grid generation procedure that defines a patched multiple-block grid system suitable for fighter-type aircraft geometries with fuselage and engine inlet, canard or horizontal tail, cranked delta wing and vertical fin has been developed. The grid generation is based on transfinite interpolation and requires little computational power. A finite-volume Euler solver using explicit Runge-Kutta time-stepping has been adapted to this grid system and implemented on the VPS-32 vector processor with a high degree of vectorization. Grids are presented for an experimental aircraft with fuselage, canard, 70-20-cranked wing, and vertical fin. Computed inviscid compressible flow solutions are presented for Mach 2 at 3.79, 7 and 10 deg angles of attack. Conmparisons of the 3.79 deg computed solutions are made with available full-potential flow and Euler flow solutions on the same configuration but with another grid system. The occurrence of an unsteady solution in the 10 deg angle of attack case is discussed.

Identification of the Longitudional Motion of a DORNIER DO 28 Airplane

The model of the propeller–driven aircraft DORNIER DO 28 of the TU Braunschweig is being identified during the last years. This nonlinear, multi–point aircraft model considers the special effects which arise from the aircraft geometry and the propeller influence. Flight test data are used for identification of the physical, aerodynamic parameters, after having been checked for data compatibility. The resulting parameters show a good consistency for well prepared flight tests.

Euler solutions for aircraft configurations employing upper-surface blowing

A method has been developed for calculating the flowfield around complex aircraft configurations using upper-surface blowing (USB). The method is based on a zonal grid generation approach coupled with an Euler flow solver algorithm. In this method the flowfield is divided into a number of non-overlapped regions, each containing a compon- ent of the aircraft such as a wing, a fuselage or a nacelle. H-type grids are generated independe- ntly in each region using a hybrid elliptic/ algebraic grid generation scheme. An explicit finite volume Euler algorithm has been developed that is applicable to the multi-region H-type grids. The present method has been applied to a two-dimensional USB model and a realistic aircraft configuration consisting of a wing/ fuselage with two integrated nacelles. Numerical results indicate that the computational method is effective in predicting the flowfield about USB-type configurations and complex aircraft geometries.

Grid generation and flow calculations for aircraft geometries

A method for calculating the flowfield around complex aircraft configurations based on a multiblock gridgeneration approach coupled with an Euler flow algorithm is presented. In this approach the flowfield is subdivided into a set of nonoverlapping blocks. Grids are generated simultaneously in all of the blocks using an elliptic grid-generation method. Appropriate boundary conditions on the block faces ensure that grid lines pass smoothly between adjacent blocks leaving a grid that is globally smooth except for a few isolated geometric singularities. A geometry package for use with the grid generator based on bicubic surface patches and with component intersection capability is briefly described and the generation of grids on the configuration surfaces is treated in some detail. An explicit finite volume Euler algorithm has been developed for use with the multiblock grids and is described herein. Some preliminary results obtained using the multiblock system are presented.

Canard-Wing Shape Optimization with Aerodynamic Requirements

THIS paper demonstrates a numerical technique for canard-wing shape optimization at two operating conditions. For purposes of simplicity, a mean surface wing paneling code1 is employed for the aerodynamic calculations. The optimization procedures2 are based on the method of feasible directions. The shape functions for describing thickness, camber, and twist are based on polynomial representations. The primary design requirements imposed restrictions on the canard and wing volumes and on the lift coefficients at the operating conditions. Results indicate that significant improvements in minimum drag and lift-to-drag ratio are possible with reasonable aircraft geometries. Calculations were done for supersonic speeds with Mach numbers ranging from 1 to 6. Planforms were mainly of a delta shape with aspect ratio of 1, with the canard and wing in the same plane. Contents The shape functions for the thickness,3 and camber, and twist4 were each expressed as a ten-term polynomial function of the Cartesian coordinates defined in the canard-wing plane. The coefficients of these polynomials had the status of optimization variables. Volumes of the wing and canard are constrained to specified values and correspond to the volumes of the base configuration in which both the canard and wing have 5% parabolic sections. The study initially explored minimizing wave drag through wing-canard shaping by calculating the optimum thickness distribution with zero camber and twist. The results are shown in Fig. 1 for two canard sizes as well as for a wing-along case. Wave drag reductions of up to 50%, relative to the base configuration with constant thickness ratio airfoils, are feasible while still meeting canard-wing internal volume limits. The improvements in drag become more pronounced at high Mach numbers. The optimum shapes were found to be similar to those reported by Strand 3 for the wing-alone case, indicating that the presence of the canard does not introduce significant perturbations in the shape functions. The second study explored the reduction in drag due to lift through optimization of the camber and twist of the lifting surface with zero thickness (Fig. 2). Again, results are shown for two canard sizes as well as for a wing-alone case. The configurations with subsonic leading edges show drag reductions of up to 36% by use of optimum camber and twist of the lifting surface. The potential for improvement tends to diminish with increasing Mach number in contrast with the results for the optimization of thickness. Figure 3 shows, in terms of LID, the data of Fig. 2 and includes a simple flat

Bizjet—A fortran program for sizing subsonic business jets

BIZJET is an iterative design algorithm coded in FORTRAN which sizes subsonic business jets. Starting from a prescribed aircraft geometry and a given jet engine. BIZJET computes the take-off gross weight, the wing planform area. and the engine size factor such that a prescribed mission can be accomplished. In addition to the sizing mode. BIZJET can be used to compute the performance of a given aircraft, or any module such as the aerodynamics module can be used independently. After a brief description of the program, examples of its use in each of the above modes are presented.

The defence review

THE authoritative leak which is today used so frequently to lessen the impact and perhaps the resistance to new statements of policy was part of the prelude to the publication of the Defence White Paper. This, together with the resignation of Mr Christopher Mayhew, the Navy Minister, three days before the publication date, meant that the main decisions—briefly, fifty F‐111A aircraft for the R.A.F., no new carriers for the Navy, a cut in the naval Phantom order, and continued support for the Anglo‐French variable geometry aircraft due to enter service in 1975—were all common knowledge beforehand and the White Paper was a mere confirmation of the expected.

An analysis of corona-generated interference in aircraft

Triboelectric charging, occurring when an aircraft is operated in precipitation, raises the aircraft potential until corona discharges occur from points of high dc field on the aircraft. These corona discharges generate noise which is coupled into receiving systems. The magnitude and spectral distribution of this radio interference, called precipitation static, depends upon three factors: 1) the strength and spectral characteristics of the source discharges, 2) the manner in which the disturbances produced by the discharges couple into the antennas, and 3) the magnitude of the discharge current and its distribution among the discharging extremities. The coupling between the antenna and the noise source is discussed using a reciprocity relationship. Because the geometry of an aircraft is complicated, and a purely theoretical approach to the determination of coupling factors is not possible, a technique developed for measuring absolute values of coupling factor as a function of frequency and position on the aircraft is described. The spectral character of the corona-noise source is studied, including a study of how the source spectrum is affected by altitude. To test the validity of the theory and the results of the laboratory work, calculations are made to predict the noise currents induced in the two test antennas employed in a flight-test program conducted on the Boeing 367-80 aircraft (prototype of the KC-135 and 707). The results of these predictions are compared with the noise spectra measured in flight.