Fuel requirement is one of the most important issues in aircraft performance. For an aircraft of a given mass, flying in an atmosphere where temperature is a monotonically decreasing function of pressure, there is one Mach number and flight level (FL) pair for which the specific air range (SAR), i.e. the distance travelled through the air per unit mass of fuel consumed, is a maximum. Specific air range is directly proportional to the range parameter, which is the product of the overall engine efficiency and the airframe lift-to-drag ratio. During the aircraft design process, the optimum value of the range parameter is fixed at a specific ‘design’ mass in a specific ‘design’ atmosphere. This ‘design optimum’ condition is a characteristic of the aircraft and engine combination, occurring at a particular Mach number and flight level pair. However, in routine operations, the aircraft mass and the atmospheric characteristics will rarely, if ever, coincide with the design conditions. Consequently, in general, the operational optimum Mach number and flight level pair will not be the same as that at the design optimum. The primary purpose of this analysis is to find the general relationship between the operational and the design conditions. In performance calculations, it is often assumed that, if the Mach number and lift coefficient are constant, the range parameter is also constant. This leads to the well-known ‘cruise climb’ trajectory for ‘maximum’ range and, for the whole flight, the classical Breguet range equation for the required trip fuel. However, to maintain a constant lift coefficient, the aircraft must climb as mass reduces and, depending upon the relationship between temperature and pressure in the atmosphere, its Reynolds number changes. It is shown that these Reynolds number changes can alter the range parameter by several percent as the flight progresses, and the standard Breguet equation may produce significant errors in the trip fuel estimation. Therefore, a completely general formulation of the range equation is presented. The well-documented, open-source, Poll-Schumann aircraft performance model is used to provide data for a wide range of operating conditions. Tabulated coefficients are provided to allow operational estimates to be made for 71 aircraft types and variants. These values are best estimates based upon information available in the public domain.

To obtain the optimal solution for the performance of the turbofan engine using infrared stealth technology, an engine mathematical model with a backward infrared radiation intensity calculation module was established. The effects of infrared suppression measures on the performance of turbofan engines were analysed. Based on the multi-objective particle swarm optimisation (MOPSO) algorithm, the optimal solution for the performance in the cruise state of the reference engine refitted with the infrared radiation suppression module was obtained; Further, through the multiple design points (MDPs) concept, the thermal cycle optimisation design of the turbofan engine was carried out. The results show that the integrated fully shielded guiding strut (IFSGS) with air film cooling had the ideal infrared suppression effect. Compared with the reference engine refitted with infrared radiation suppression module, the engine after cycle optimisation design could obtain better infrared stealth performance.

The aviation industry relies on accurate models. These models are used to predict an aircraft system’s outputs, and thus allow an understanding of the parameters involved, which could lead to system improvements. This study focuses on the engine modelling of an aircraft, and on its experimental validation using the Cessna Citation X Research Aircraft Simulator designed by CAE Inc., equipped with a level D Flight Dynamics toolbox. Level D is the highest rank attributed by the Federal Aviation Administration FAA certification authorities for flight dynamics. The proposed model aims to predict the thrust and the fuel consumption for various altitudes, Mach numbers and throttle lever angles (TLA). Different generic static models, which correspond to their steady state, from the literature, were used in this study; however, most of them were validated under restricted hypotheses. An optimisation algorithm was used in order to tune the static model parameters with the set of identification flight test data. Another set of data was then used in order to validate the identified model. Furthermore, a dynamic model corresponding to the transient operations was identified. TLA steps, impulses and ramp perturbations were performed in order to identify the system response, and to validate system dynamic model with other flight tests than the identification tests.

The next generation of civil large aero-engines will employ greater bypass ratios compared with contemporary architectures. This results in higher exchange rates between exhaust performance and specific fuel consumption (SFC). Concurrently, the aerodynamic design of the exhaust is expected to play a key role in the success of future turbofans. This paper presents the development of a computational framework for the aerodynamic design of separate-jet exhaust systems for civil aero-engines. A mathematical approach is synthesised based on class-shape transformation (CST) functions for the parametric geometry definition of gas-turbine exhaust components such as annular ducts and nozzles. This geometry formulation is coupled with an automated viscous and compressible flow solution method and a cost-effective design space exploration (DSE) approach. The framework is deployed to optimise the performance of a separate-jet exhaust for very-high-bypass ratio (VHBR) turbofan engine. The optimisations carried out suggest the potential to increase the engine’s net propulsive force compared with a baseline architecture, through optimum exhaust re-design. The proposed method is able to identify and alleviate adverse flow-features that may deteriorate the aerodynamic behaviour of the exhaust system.