Authors:D.I.A. Poll; U. Schumann Pages: 257 - 295 Abstract: This paper is one of a series addressing the need for simple, yet accurate, methods for the estimation of cruise fuel burn and other important aircraft performance parameters. Here, a previously published, constant Reynolds number model for turbofan-powered, civil transport aircraft is extended to include Reynolds number effects. Provided the variation of temperature with pressure is known, the method is applicable to flight in any atmospheric conditions. For a given aircraft, cruising in a given atmosphere, there is a single Mach number and Flight Level pair, at which the fuel burn per unit distance travelled through the air has an absolute minimum value. Both these quantities depend upon the Reynolds number, which, in turn, depends upon the aircraft weight and the atmospheric vertical temperature profile. Simple, explicit expressions are developed for all parameters at the optimum condition. These are shown to be in close agreement with numerical solutions of the governing equations. It is found that typical operational mass and temperature profile variations can change cruise fuel burn rate by several percent. In the International Standard Atmosphere, when the speed and altitude deviate from their optimum values, the fuel burn penalty is reduced slightly relative to the constant Reynolds number case. By way of example, the method is used to estimate the minimum fuel, speed-versus-height trajectory for cruise in a realistic atmosphere.For each aircraft, cruise fuel burn is found to be governed by six independent parameters. All are constants. Two are simple, involving only size and weight, whereas four are complex and must be determined by either theoretical, or empirical, means. The estimation of these quantities will be considered in Part 2. PubDate: 2021-02-01T00:00:00.000Z DOI: 10.1017/aer.2020.62 Issue No:Vol. 125, No. 1284 (2021)
Authors:D.I.A. Poll; U. Schumann Pages: 296 - 340 Abstract: A simple yet physically comprehensive and accurate method for the estimation of the cruise fuel burn rate of turbofan powered transport aircraft operating in a general atmosphere was developed in part 1. The method is built on previously published work showing that suitable normalisation reduces the governing relations to a set of near-universal curves. However, to apply the method to a specific aircraft, values must be assigned to six independent parameters and the more accurate these values are the more accurate the estimates will be. Unfortunately, some of these parameters rarely appear in the public domain. Consequently, a scheme for their estimation is developed herein using basic aerodynamic theory and data correlations. In addition, the basic method is extended to provide estimates for cruise lift-to-drag ratio, engine thrust and engine overall efficiency. This step requires the introduction of two more independent parameters, increasing the total number from six to eight. An error estimate and sensitivity analysis indicates that, in the aircraft’s normal operating range and using the present results, estimates of fuel burn rate are expected to be in error by no more than 5% in the majority of cases. Initial estimates of the characteristic parameters have been generated for 53 aircraft types and engine combinations and a table is provided. PubDate: 2021-02-01T00:00:00.000Z DOI: 10.1017/aer.2020.124 Issue No:Vol. 125, No. 1284 (2021)
Authors:A. Khalil; N. Fezans Pages: 341 - 364 Abstract: Turbulence and gusts cause variations in the aerodynamic forces and moments applied to the structure of aircraft, resulting in passenger discomfort and dynamic loads on the structure that it must be designed to support. By designing Gust Load Alleviation (GLA) systems, two objectives can be achieved: first, realizing higher passenger comfort; and second, reducing the dynamic structural loads, which allows the design of lighter structures. In this paper, a methodology for designing combined feedback/feedforward GLA systems is proposed. The methodology relies on the availability of a wind profile ahead of the aircraft measured by a Doppler LIDAR sensor, and is based on $H_{\infty}$-optimal control techniques and a discrete-time preview-control problem formulation. Moreover, to allow design trade-offs between those two objectives (to achieve design flexibility) as well as to allow specification of robustness criteria, a variant of the problem using multi-channel $H_{\infty}$-optimal control techniques is introduced. The methodology developed in this paper is intended to be applied to large aircraft, e.g. transport aircraft or business jets. The simulation results show the effectiveness of the proposed design methodology in accounting for the measured wind profile to achieve the two mentioned objectives, while ensuring both design flexibility and controller robustness and optimality. PubDate: 2021-02-01T00:00:00.000Z DOI: 10.1017/aer.2020.85 Issue No:Vol. 125, No. 1284 (2021)
Authors:J. Hollom; N. Qin Pages: 365 - 388 Abstract: Uncertainty in the critical amplification factor ($N_{cr}$) of the $e^N$ transition model is used to approximate the uncertainty in the surface and flow quality of natural laminar flow (NLF) aerofoils. The uncertainty in $N_{cr}$ is represented by a negative half-normal probability distribution that descends from the largest $N_{cr}$ achievable with an ideal surface and flow quality. The uncertainty in various aerodynamic coefficients due to the uncertainty in $N_{cr}$ is quantified using the weighted mean and standard deviation of flow solutions run at different $N_{cr}$ values. The uncertainty in the aerofoil performance is assessed using this methodology. It is found that the standard deviation of the aerofoil performance due to the uncertainty in $N_{cr}$ is largest when the transition location is most sensitive to changes in the lift coefficient at the ideal $N_{cr}$. Robust shape optimisation is also carried out to improve the mean performance and reduce the standard deviation of the performance with uncertainty in $N_{cr}$. This is found to be effective at producing aerofoils with a larger amount of laminar flow that are less sensitivity to uncertainty in $N_{cr}$. A trade-off is observed between the mean performance and the standard deviation of the performance. It is also found that reducing the standard deviation of the performance at one Mach n... PubDate: 2021-02-01T00:00:00.000Z DOI: 10.1017/aer.2020.63 Issue No:Vol. 125, No. 1284 (2021)
Authors:H.X. Xiong; S.H. Yi, H.L. Ding, L. Jin, J.J. Huo Pages: 389 - 409 Abstract: In the development process of high-speed aircraft, the head of the aircraft is subject to high temperatures and high speed flows, supporting the maximum heat flow and thus requiring a reliable cooling system. A new type of head cooling system is proposed herein. An internal flow channel model of the heat transfer in a ball head made from high-temperature alloy steel is constructed, then an experimental platform is built to carry out relevant experiments on the performance of this cooling system. Firstly, the influence of different experimental conditions on the cooling efficiency of the ball head is studied. For given liquid-nitrogen supply pressure, a higher heating heat flux density on the outer surface of the ball head corresponds to higher cooling efficiency. Then, the vaporisation effect under different experimental conditions is evaluated using temperature sensors at the inlet and outlet of the ball head heat exchange channel in combination with images of the visualised glass tube. It is found that liquid nitrogen can vaporise completely when flowing through the heat exchange channel. The characteristics of the heating effect and liquid nitrogen injection for the ball head were evaluated using an infrared camera. Finally, under different experimental conditions of liquid-nitrogen supply pressure, it is found that liquid nitrogen can vaporise completely in each case, and the total temperature of the vaporised nitrogen is about 300K. It can thus be collected as a secondary gas source. PubDate: 2021-02-01T00:00:00.000Z DOI: 10.1017/aer.2020.86 Issue No:Vol. 125, No. 1284 (2021)
Authors:R.F. Latif; M.K.A. Khan, A. Javed, S.I.A. Shah, S.T.I. Rizvi Pages: 410 - 429 Abstract: We present a hybrid, semi-analytical approach to perform an eigenvalue-based flutter analysis of an Unmanned Aerial Vehicle (UAV) wing. The wing has a modern design that integrates metal and composite structures. The stiffness and natural frequency of the wing are calculated using a Finite Element (FE) model. The modal parameters are extracted by applying a recursive technique to the Lanczos method in the FE model. Subsequently, the modal parameters are used to evaluate the flutter boundaries in an analytical model based on the p-method. Two-degree-of-freedom bending and torsional flutter equations derived using Lagrange’s principle are transformed into an eigenvalue problem. The eigenvalue framework is used to evaluate the stability characteristics of the wing under various flight conditions. An extension of this eigenvalue framework is applied to determine the stability boundaries and corresponding critical flutter parameters at a range of altitudes. The stability characteristics and critical flutter speeds are also evaluated through computational analysis of a reduced-order model of the wing in NX Nastran using the k- and pk-methods. The results of the analytical and computational methods are found to show good agreement with each other. A parametric study is also carried out to analyse the effects of the structural member thickness on the wing flutter speeds. The results suggest that changing the spar thickness contributes most significantly to the flutter speeds, whereas increasing the rib thickness decreases the flutter speed at high thickness values. PubDate: 2021-02-01T00:00:00.000Z DOI: 10.1017/aer.2020.71 Issue No:Vol. 125, No. 1284 (2021)
Authors:T. Lin; W. Xia, S. Hu Pages: 430 - 451 Abstract: Lack of flexibility limits the performance enhancement of man-made flapping wing Micro Air Vehicles (MAVs). Active chordwise deformation (bending) is introduced into the flapping wing model at low Reynolds number of Re = 200 in the present study. The lattice Boltzmann method with immersed boundary is adopted in the numerical simulation. The effects of the bending amplitude, bending frequency and phase lag between bending and flapping on the propulsive performance are analysed. The numerical results show that all the chordwise deformation parameters including the bending amplitude, bending frequency and phase lag have a great influence on the flow field, Leading-Edge Vortex (LEV), Trailing-Edge Vortex (TEV) and previous Leading-Edge Vortex (pLEV) of the deformable flapping wing, which leads to the variation of the propulsive performance. With decreasing bending amplitude and increasing bending frequency, both the thrust and energy dissipation coefficients increase. The highest thrust coefficient and highest energy dissipation coefficient occur at a phase lag of 180°. On the other hand, strong dependence of the propulsive efficiency on the vortex tangle is found. The highest propulsive efficiency is obtained for the present model at a dimensionless bending amplitude of 0.2, bending frequency of 0.7Hz, and phase lag of 0°. PubDate: 2021-02-01T00:00:00.000Z DOI: 10.1017/aer.2020.72 Issue No:Vol. 125, No. 1284 (2021)
Hybrid journal (It can contain Open Access articles)
$H_{\infty}$-optimal control techniques and a discrete-time preview-control problem formulation. Moreover, to allow design trade-offs between those two objectives (to achieve design flexibility) as well as to allow specification of robustness criteria, a variant of the problem using multi-channel 
$H_{\infty}$-optimal control techniques is introduced. The methodology developed in this paper is intended to be applied to large aircraft, e.g. transport aircraft or business jets. The simulation results show the effectiveness of the proposed design methodology in accounting for the measured wind profile to achieve the two mentioned objectives, while ensuring both design flexibility and controller robustness and optimality.
$N_{cr}$) of the 
$e^N$ transition model is used to approximate the uncertainty in the surface and flow quality of natural laminar flow (NLF) aerofoils. The uncertainty in 
$N_{cr}$ is represented by a negative half-normal probability distribution that descends from the largest 
$N_{cr}$ achievable with an ideal surface and flow quality. The uncertainty in various aerodynamic coefficients due to the uncertainty in 
$N_{cr}$ is quantified using the weighted mean and standard deviation of flow solutions run at different 
$N_{cr}$ values. The uncertainty in the aerofoil performance is assessed using this methodology. It is found that the standard deviation of the aerofoil performance due to the uncertainty in 
$N_{cr}$ is largest when the transition location is most sensitive to changes in the lift coefficient at the ideal 
$N_{cr}$. Robust shape optimisation is also carried out to improve the mean performance and reduce the standard deviation of the performance with uncertainty in 
$N_{cr}$. This is found to be effective at producing aerofoils with a larger amount of laminar flow that are less sensitivity to uncertainty in 
$N_{cr}$. A trade-off is observed between the mean performance and the standard deviation of the performance. It is also found that reducing the standard deviation of the performance at one Mach n...
