A numerical design synthesis methodology for new generations of combat aircraft has been developed. It incorporates advanced technology in the form of design for low observables. Aircraft capable of being modelled with this methodology will have internal or external Weapons carriage, side mounted intakes, a straight-tapered trapezoidal wing, aft-mounted tail with the option of single or twin ns, and one or two engines with rectangular or axisymmetric nozzles. The design methodology incorporates sufficiently accurate and realistic algorithms for the calculation of the geometry and the estimation of the aerodynamic, mass and performance properties of the aircraft. The inherent flexibility of the design permits the examination of a wide range of configurations whilst maintaining the accuracy required to examine minor changes in the design requirements. A numerical optimization routine was linked to the synthesis, allowing the determination of optimum aircraft design variables for a given set of mission and performance requirements. Results were obtained showing the usefulness of this design tool for setting up parametric trend studies. The numerical accuracy, flexibility of configuration options and high level of advanced aircraft technology of this synthesis make a significant contribution to the continuing development of automated design tools.

A technique for the identification of aircraft stability and control parameters from flight test recordings made in either calm or turbulent air is presented. The maximum likelihood output error method is used with a steady-state Kalman filter incorporated to account for atmospheric turbulence. A modified Newton-Raphson search technique, enhanced by a line search, is employed for parameter identification. Separate algorithms are developed for estimating the biases and noise levels in the observations. Some areas of practical problems in the application of such methods are stressed. A computer program for the identification of longitudinal stability derivatives is described and the aircraft instrumentation required is exemplified by that in Gnat XPSOS. The wind tunnel calibration of the flow-direction sensing nose probe assembly of Conrad yawmeters on this aircraft is detailed. The problems of handling the flight observations recorded are covered and some of the troubles experienced with the instruments are noted. The performance of the identification technique is investigated. The data required, the choices open to the analyst and the statistical information produced being highlighted. The identification of the process noise level, in this instance the turbulence intensity, is addressed and it is found that the value specified for this level can influence the other parameters. The longitudinal stability and control derivatives obtained for Gnat XPSOS are presented.

Generally speaking, difficulties encountered during the improvement of an Analytical Dynamic Model (ADM) using vibration test data come from both the Spatial Coordinate Incompatibility (SCI) and especially the Modal Coordinate Incompatibility (MCI) between the ADM and the test data. Efforts were therefore made in this project to cope with these two problems by extending some of the existing methods and also by developing new methods with consideration of their feasibility, efficiency and accuracy. A general description of this part of the project and the literature survey of this study area are presented in part 1 of the thesis. In part 2, in order to solve the SCI problem, a new extended Complete Modal Expansion (CME) and a Branch Modal Expansion (BME) method were proposed especially for the case when using a branch mode method to produce the ADM. Application of these two methods and the existing physical expansion method were demonstrated in a beam example in this part and were also used in some of the examples later in this project. In part 3, efforts were made to extend the existing Direct Matrix Updating (DMU) and the Direct Parameter Identification (DPI) methods for solving the MCI problem using a direct approach. Firstly a new Direct Modal Extension (DME) method was proposed and compared with the DMU method when they were used to improve a reduced-size ADM. Secondly, in order to overcome the main limitation of the existing DPI methods in their practical use, an extended Corrected Modal Constraint (CMC) method was proposed. In part 4, in order to achieve the feasibility and accuracy of ADM improvement, efforts were then made in the study of the indirect approach. Firstly a procedure using a new Orthogonality Sensitivity Method (OSM) working together with a model reduction method was proposed. Secondly, a new Energy Error Estimation (EEE) method was also presented. The original contribution of the EEE method is that the poorly modelled stiffness and mass elements of an ADM can be identified and corrected accurately and effectively. Applications of these new proposed methods were demonstrated by taking beam examples. Further application of the EEE method was examined in a full-scale aircraft tail plane example. A general discussion, conclusion and recommendation for further study of these methods are presented in the fmal part 5. Based on the study of this project, it is concluded that the feasibility and accuracy of the direct methods described in part 3 of this thesis are at a low level for practical use. Therefore, the main efforts and contributions in this project were made in the study of the indirect methods described in part 4 of this thesis. It is concluded that both of the new proposed OSM and EEE methods provide feasible tools for ADM improvement and possess a high level of accuracy.

The purpose of the research investigation was to study the application of semi-active control technology to the design of a suspension system to be used in a landing gear of a high speed military aircraft. A semi-active system was used because it will allow a system to be driven from the hydraulic systems already existing in the aircraft without extensive modification. The research work involved establishing a theoretical mathematical model for the semi-active damping system. This model involved a large number of non-linear dynamic phenomena and elements including a two-stage gas spring, lever geometry, break out friction, square law damping and the switching function needed to achieve the semi-active control. Validation of the model was carried out by means of an extensive study of the dynamic responses obtained from digital simulation. An extended programme of laboratory experiments was also carried out to confirm the theoretical and simulated results, and to demonstrate the potential benefits in performance which can be achieved with those obtained from standard and optimized passive suspension system. The experimental rig involved a physical model which used hydraulic elements of a general industry standard, but not specially approved for aircraft use. The apparatus was arranged to permit a considerable degree of freedom for implementing the control laws which facilitated the assessment of different control schemes and allowed, at the same time, the ready simulation of various passive damping arrangements. An extensive series of trials was carried out on the final design and involved frequency response tests and subjecting the experimental suspension to inputs obtained from a simulated runway profile. The profile simulation was a discrete representation of a particular runway chosen for its roughness which was characteristic of runways from which high-speed military aircraft operate. From the research investigation and these trials it was established that semi-active control of the damping function is superior to standard techniques and achieves a substantial reduction in the energy transmitted to the airframe during ground manoeuvres.

No description available.

No description available.

Designing successful commercial aircraft is a difficult business; the stakes are high and the risks numerous. Researchers in the past have developed methods that assist the designers in reducing these risks. In recent years such methods have benefited from improvements in computer technology. The work described in this thesis extends these methods to the design of commuter aircraft. These aircraft are more sensitive to operational requirements than other types due in part to their high zero-fuel mass ratio. It is essential that, for such aircraft, the best information possible is available to the designers. The identification of the optimum aircraft configuration and mission characteristics constitutes a vital part of this knowledge. A review of literature, involving both modem computer-based and traditional search methods, has shown continuing interest in aircraft project design methods from the earliest times to the latest conference. The work presented in this thesis is seen to compliment this interest in computer methods and to apply these techniques to the relatively neglected area of commuter aircraft design. A survey of commuter operation and aircraft types revealed the often conflicting requirements and regulations which govern the design process in this area. Detailed statistical analysis on a collection of commuter aircraft showed no consistent data patterns, but did indicate the bouyant state of the market. Earlier research work on the design of twin-engined turbo-prop aircraft had provided some experience in the design of short-haul aircraft. The new work improves these methods and applies them to larger and faster turbo-fan commuter aircraft. Since the turbo-prop work, the optimiser developed at RAE (Farnborough) has been rewritten to work more efficiently and allow larger problems to be tackled. This new optimiser s linked to a new synthesis routine which simulates turbo-fan aircraft design. The synthesis program was calibrated against industrial design calculations and shown to give acceptably accuracte predictions. The resulting design program is fully described and computer listings are presented. To illustrate the use of the optimisation methods in the devleopment of a new aircraft, a series of industrially related design studies is presented. These studies range from the selection of the initial baseline configuration, through various parameters sensitivity investigations, to the evaluation of aircraft and engine stretch options. To demonstrate more general types of design study, a series of optimisations in which the engine size is variable was conducted. This provides the designer with a knowledge of the absolute (optimum) design surface and allows him to judge the 'penalties' inherent in his chosen configuration.

No description available.

Vertical/Short Take-Off arid Landing (V/STOL) fighter aircraft are characterised by increased control complexity caused by the extra degree ol freedom. This can result in a high pilot workload which may be alleviated with the careful application of active flight control. However, the advent of control configured vehicles demands that the controller design must be part of a fully integrated and iterative aircraft design; hence it must allow the two-way flow of design information. In this thesis a suitable controller design method is developed to solve this two-fold problem. The method is based upon a singular perturbation analysis which is used to expose the underlying dynamics of a closed-loop state-space system. developments are described which allow high-order, dynamically complex parasitics, such as actuators, to be included in the design. Furthermore, the method gives the designer insight into the problem allowing tuning and engineering trade-offs to be performed intelligently with a two-way flow of design information. The end result is a robust high-gain multivariable controller. In order fully to develop arid analyse the method it has been applied to a representative non-linear time-varying aircraft simulation model. This LS supplied by the Royal Aerospace Establishment, Bedford. The necessary slate-space matrices are otitairted by lirLearisirig the model at several different flight cases. This occurs over a wide flight envelope, from hover to 300 Kts, and consequently the multivariable control laws are implemented using gain scheduling. Finally, task tailored control and handling qualities requirements are derived for a V/STOL aircraft in the form of a design brief. This design brief is then fulfilled by designing a controller which alleviates pilot workload during transitions from jet-borne to fully wing-borne flight (and vice versa).

No description available.