Design Of Aircraft Configurations Research Articles

Multidisciplinary Integrated Preliminary Design Applied to Unconventional Aircraft Configurations

A preliminary aircraft design tool is presented as a means of performing multidisciplinary, integrated preliminary design of unconventional aircraft configurations. Higher fidelity numerical methods for some of the involved disciplines are discussed as key elements of this process. A noise propagation module is described as a first step toward introducing aircraft noise analysis into the preliminary design process. Application of the tool to a specific unconventional aircraft concept is shown. The results of this study provide an understanding of why multidisciplinary design analysis and optimization, as well as higher fidelity methods, are required for the design of unconventional aircraft configurations.

항공기의 실속 회복을 위한 자동 회복 장치 설계 및 검증에 관한 연구

Modem version of supersonic jet fighter aircraft must have been guaranteed appropriate controllability and stability in HAoA(High Angle of Attack). Limit value of aircraft entering into the departure in HAoA is related to aircraft configuration design. But, the control law such as AoA and yaw-rate limiter is implemented in digital Fly-By-Wire flight control system of supersonic jet fighter to guarantee the aircraft`s safety in HAoA. The HAoA flight control law have two parts, one is control law of departure prevention and the other is control law of departure recovery support. The control laws of departure prevention for advanced jet trainer consist AoA limiter, roll command limiter and rudder fader. The control laws of departure recovery support are consist yaw-rate limiter and MPO(Manual Pitch Override) mode. The guideline of pitch rocking using MPO mode is simple, but operating skill of pitch rocking is very difficult by the pilot with inexperience of departure situation. This paper addresses the design and validation of APR(Automatic Pitch Rocker) control law instead of MPO in order to automatic recovery without manual pitch rocking by the pilot. And, recovery characteristic with APR verifies by the nonlinear analysis and pilot evaluation.

Multi--Disciplinary Optimization for the Conceptual Design of Innovative Aircraft Configurations

Multi--Disciplinary Optimization for the Conceptual Design of Innovative Aircraft Configurations

고받음각에서 조종성능 및 안정성 증강을 위한 제어법칙에 관한 연구

현대의 고성능 전투기는 한계받음각 내에서 적절한 조종성(Controllability) 및 안정성(Stability)을 확보하고 있어야 한다. 고받음각에서 항공기가 이탈에 진입할 수 있는 한계값은 항공기 형상설계에 직결되는 문제이다. 하지만 현대의 고성능 전투기는 전기식 비행제어계통(Digital Fly-By-Wire Flight Control System)을 사용하여 고받음각 제어법칙을 설계함으로써 한계값 내에서 항공기의 안정성을 보장하고 있다. 현재, T-50 세로축 고받음각 제어법칙에는 (+)한계받음각 이상으로의 비행을 제한하는 받음각 제한기(Angle of Attack Limiter)가 설계되어 있다. 그러나 (-)한계받음각 이상으로 비행을 제한하는 받음각 제한기는 설계되어 있지 않아, 조종사의 과도한 (-)세로축 기동에 대해 항공기가 이탈에 진입할 수 있다. 그리고 T-50 비행시험에서 PA에서 과도한 (+)세로축 기동 시, 받음각(Angle of Attack)이 현재 설계되어 있는 한계받음각을 초과하는 문제점이 발생하였다. 본 논문에서는 고받음각 제어법칙을 일부 수정하여 한계받음각 내에서 항공기의 세로축 조종성능을 향상시키고, 항공기 안정성을 확보할 수 있는 제어법칙에 관한 연구를 수행하였다. Modern version of supersonic jet fighter aircraft must have guaranteed appropriate controllability and stability in HAoA(high angle of attack). Limit value of aircraft entering into the deep stall in HAoA is related to aircraft configuration design. But, In order to guarantee the aircraft's safety in HAoA, control law for HAoA region implemented in digital Fly-By-Wire flight control system of supersonic jet fighter. The AoA limiter is designed for positive HAoA in longitudinal control law. But, aircraft departure during aggressive negative pitch maneuver such as push over in departure resistance flight test. Therefore negative AoA limiter is needed in longitudinal control law. Result of T-50 flight test show that the AoA is exceed the limit value during aggressive positive pitch maneuver in pull up of power approach mode. In this paper, the AoA limit control law in positive and negative AoA was proposed in order to improve aircraft controllability and stability.

고받음각에서 항공기 이탈 방지를 위한 제어법칙에 관한 연구

군용항공기는 고받음각에서 적절한 조종성 및 항공기 이탈에 대한 안정성을 확보하고 있어야 한다. 고받음각에서 항공기가 이탈에 진입할 수 있는 한계값은 항공기 형상설계에 직결되는 문제이다. 하지만 현대의 고성능 전투기는 전기식 비행제어계통을 사용하여 고받음각 제어법칙을 설계함으로써 한계 값 내에서 항공기의 안정성을 보장하고 있으며 항공기가 이탈 시, 안전하게 회복할 수 있도록 고받음각 제어법칙을 설계한다. 현재 T-50에 적용되어 있는 고받음각 제어법칙은, 세로축 방향으로는 받음각 제한기 및 MPO(Manual Pitch Override) 모드, 가로-방향축으로는 고받음각 이탈제한기, 가로축 명령제한기, 방향축 조종사 명령제한기 및 스핀방지기가 설계되어 있다. 본 논문에서는 T-50 에 적용되어 있는 고받음각 제어법칙을 소개하며, 고받음각 비행시험을 통하여 항공기 안정성에 관한 연구를 수행하였다. Supersonic jet fighter aircraft must have been guaranteed appropriate for controllability and stability in HAoA(High Angle of Attack) region. Limit value of aircraft enter the deep stall at HAoA is related to problem of aircraft configuration design. But, In order to guarantee the aircraft safety in HAoA, control law is designed using digital Fly-By-Wire flight control system in modern versions of supersonic jet fighter aircraft. Also, In order to recovery if aircraft enter the deep stall or spin, anti-spin control law and MPO(Manual Pitch Override) mode is designed. AoA limiter and MPO is designed in longitudinal axis and HAoA departure prevention logic, roll command limiter, rudder fader and anti-spin logic is designed in lateral-directional axis. In this paper, we introduce the T-50 HAoA flight control law and propose that aircraft stability and adequate of these control law from HAoA flight test.

VizCraft: a problem-solving environment for aircraft configuration design

The VizCraft problem-solving environment aids aircraft designers during conceptual design of a high-speed civil transport (HSCT). It integrates simulation codes that evaluate a design with visualizations for analyzing a design individually or in contrast to other designs. VizCraft provides a graphical user interface to a widely used suite of simulation and analysis codes for HSCT design, and it provides tools for visualizing the outputs of these codes. So, VizCraft provides an environment that combines visualization and computation, encouraging the designer to think in terms of the overall problem-solving task, not simply using the visualization to view the computation's results.

Constrained Multipoint Aerodynamic Shape Optimization Using an Adjoint Formulation and Parallel Computers, Part 1

An aerodynamic shape optimization method that treats the design of complex aircraft configurations subject to high fidelity computational fluid dynamics (CFD), geometric constraints and multiple design points is described. The design process will be greatly accelerated through the use of both control theory and distributed memory computer architectures. 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 implemented on parallel distributed memory architectures using a domain decomposition approach, an optimized communication schedule, and the MPI (Message Passing Interface) standard for portability and efficiency. The final result achieves very rapid aerodynamic design based on a higher order CFD method. In order to facilitate the integration of these high fidelity CFD approaches into future multi-disciplinary optimization (NW) applications, new methods must be developed which are capable of simultaneously addressing complex geometries, multiple objective functions, and geometric design constraints. In our earlier studies, we coupled the adjoint based design formulations with unconstrained optimization algorithms and showed that the approach was effective for the aerodynamic design of airfoils, wings, wing-bodies, and complex aircraft configurations. In many of the results presented in these earlier works, geometric constraints were satisfied either by a projection into feasible space or by posing the design space parameterization such that it automatically satisfied constraints. Furthermore, with the exception of reference 9 where the second author initially explored the use of multipoint design in conjunction with adjoint formulations, our earlier works have focused on single point design efforts. Here we demonstrate that the same methodology may be extended to treat complete configuration designs subject to multiple design points and geometric constraints. Examples are presented for both transonic and supersonic configurations ranging from wing alone designs to complex configuration designs involving wing, fuselage, nacelles and pylons.

Sonic boom acceptability studies at NASA Langley Research Center

As part of the High Speed Research Program at NASA, Langley Research Center conducted a series of studies during the period 1990–1995 to evaluate human responses to sonic boom impact. This work, in combination with aircraft configuration design efforts, was aimed at determining the feasibility of overland supersonic flight. NASA conducted a three part program, comprising laboratory studies, in-home field experiments and community surveys. The laboratory studies used a sonic boom simulator and were designed to quantify relative loudness and annoyance response to a wide range of shaped boom signatures that might be produced by advanced aircraft configurations. The primary conclusion was that several common noise metrics were able to reliably predict judged loudness/annoyance. Field experiments were conducted in people’s homes using a noise simulation system. A controlled sound environment of booms of known level and number was introduced into a realistic home setting and human responses to a variety of boom exposures were assessed. It was found that the ‘‘equal energy’’ hypothesis was applicable to these boom exposures. Community surveys were conducted in two areas routinely impacted by sonic booms from military aircraft. Relationships between boom exposure and residents annoyance were developed for the two survey areas.

A CASCADE OPTIMIZATION STRATEGY FOR SOLUTION OF DIFFICULT DESIGN PROBLEMS

A research project to evaluate comparatively ten different non-linear optimization algorithms was completed recently. A conclusion was that no single optimizer could successfully solve all the 40 structural design problems in the test-bed, even though most optimizers successfully solved at least one-third of the problems. We realized that improvements to search directions and step lengths, available in the ten optimizers compared, were not likely to alleviate the convergence difficulties. For the solution of those difficult problems we have devised an alternate approach called, the cascade optimization strategy. The strategy utilizes several optimizers, one followed by another in a specified sequence, to solve a problem. A pseudo-random dumping scheme perturbs the design variables between the optimizers. The cascade strategy has been tested out successfully in the design of supersonic and subsonic aircraft configurations and air breathing engines for high-speed civil transport applications. These problems could not be successfully solved by an individual optimizer. The cascade optimization strategy, however, generated feasible optimum solutions for both aircraft and engine problems. This paper presents the cascade strategy, solution of aircraft and engine problems along with discussions and conclusions. © 1997 John Wiley & Sons, Ltd.

A Subsonic Panel Method for Iterative Design of Complex Aircraft Configurations

An iterative-design method for the generation of wing geometries with specified surface-pressure distributions is described. The design method is an adaptation of the well-known surface-transpiration technique for approximating boundary-layer effects. A velocity distribution normal to the aircraft surface is sought which minimizes the difference between computed and desired pressures. The transpiration distribution for a given initial wing geometry is determined by numerical optimization and is then used to reloft the configuration. The use of this iterative procedure with a subsonic surface-singularity panel method is illustrated for several example design problems. I. Introduction D URING the past decade, several panel methods16 have been developed which can be used to determine the potential flowfield about realistic aircraft shapes. These computational methods provide the needed capability for accurately modeling the aerodynamic interference effects of arbitrary geometries such as multiple lifting surface and/or wing-body configurations. Viscous effects for weak interaction flow conditions can also be approximated by the incorporation of suitable boundary-layer computational methods.1

Advances in Supersonic Configuration Design Methods

A linearized panel method program developed for the design and analysis of wing-body configurations at supersonic speeds (Woodward 1) has been modified to permit the user to specify the loading for an arbitrary region of the wing. The remaining portion of the wing is then optimized on the basis of the prescribed loading. This insures that the final shape will produce the minimum drag possible within the constraints. In addition, the design feature has been incorporated into an improved panel method supersonic analysis program (Woodward II). Since this code has the capability of representing bodies of noncircular cross section, it is expected to prove useful in the design of aircraft configurations with arbitrarily shaped fuselages. Comparisons are presented with other known minimum drag solutions which demonstrate the validity of both new techniques.

Recent applications of advanced computational methods in the aerodynamic design of transport aircraft configurations

Aerodynamic designers of transport aircraft have been dreaming for decades of being able to use, for high speed design problems, some of the advanced computational tools that they now have available to them such as the three-dimensional transonic flow methods and three-dimensional finite-difference boundary-layer methods. It is anticipated by the designer and his management that the appropriate use of these new advanced computational methods on the wing design of the next subsonic transport aircraft configuration will typically result in an improved aerodynamic technology level, an improved aerodynamic efficiency for a given level of aerodynamic technology, and reduced design costs through a reduction in the amount of wind tunnel testing required and the resulting shortened time period required to define the final lines.