Helicopter downwash impact on ship airwake is addressed from a three-pronged approach: (1) Analysis of one-point statistics of autospectrum and two-point statistics of cross-spectrum and coherence from a Computational Fluid Dynamics database of flow velocities effected by helicopter downwash and shipboard airwake; (2) Development of a mathematical framework for extracting interpretive turbulence models in closed form from these autospectral statistics; and (3) Simulation through white-noise-driven filters for the extracted models. The framework begins with an earlier-exercised perturbation-type series expansion of the autocorrelation for all three velocity components, where the first term of the series has a form of the von Karman longitudinal or lateral correlation function. After transformation into equivalent series of autospectrum, the coefficients in the series are evaluated by satisfying theoretical constraints and fitting a curve on a set of selected autospectral data points generated from the database. The framework represents a sensible combination of series expansion, exploitation of a database, and theoretical constraints to provide a foothold on airwake-downwash phenomenon for engineering analysis. It ensures that the extracted model and the autospectral data points have the same time scale, mean square value, and asymptotic decay according to the Kolmogorov –5/3 Law. The framework's strengths and weaknesses, and its major advancement over the earlier series-expansion schemes are also addressed. Finally it is shown that downwash increases airwake energy (mean square value) by one order of magnitude, and almost all of this airwake-downwash energy is concentrated within the bandwidth (0.16 < f(Hz) < 1.6) that affects flight mechanics.

The fast determination of aerodynamic parameters such as pressure distributions, lift, drag and moment coefficients from the known airflow conditions (angles of attack, Mach and Reynolds numbers) in real time is still not easily achievable by numerical analysis methods in aerodynamics and aeroelasticity. A flight parameters control system is proposed to solve this problem. This control system is based on new optimisation methodologies using Neural Networks (NNs) and Extended Great Deluge (EGD) algorithms. Validation of these new methodologies is realised by experimental tests using a wing model installed in a wind tunnel and three different transducer systems (a FlowKinetics transducer, an AEROLAB PTA transducer and multitube manometer tubes) to determine the pressure distribution. For lift, drag and moment coefficients, the results of our approach are compared to the XFoil aerodynamics software and the experimental results for different angles of attack and Mach numbers. The main purpose of this new proposed control system is to improve, in this paper, wing aerodynamic performance, and in future to apply it to improve aircraft aerodynamic performance.

Formation flight could greatly assist the air transport industry in tackling the challenges of environmental impact, excessive reliance on fuel and overcapacity. Previous studies have shown drag reductions leading to significant fuel savings for aircraft in formation relative to their solo flight. Safety is guaranteed with the use of extended formation distances, and practical implementation issues could be solved in the near future. Since studies so far have focused on existing aircraft configurations and technology, a case study using a strut-braced wing airliner was carried out to ascertain its applicability to less conventional craft. The present results did not indicate such clear-cut benefits. If formation flight is to be successful and beneficial for the next generations of aircraft, it will be vital to consider its interaction with new technologies developed for highly efficient operation, in particular those aimed at reduction of aircraft drag such as laminar flow, and to do so early in the design of aerospace vehicles and wider systems.

A method based on the Eulerian two-phase flow theory to numerically simulate three-dimensional rime ice accretions on an aircraft wing is presented in this paper. The governing equations for supercooled droplet motion under Eulerian framework are established using the droplet pseudo-fluid model. A permeable wall boundary condition is proposed to simulate the phenomenon of droplets impinging on the wing in solving the governing equations for droplets. The local droplet collection efficiency is readily obtained from the droplet flowfield solution in the control volume adjacent to the wing surface. The rime ice accretion can be simulated under the assumption that the droplets freeze immediately as they impinge on the wing surface since the environment temperature is low enough (typically below –15°C). A method to build the ice shape is proposed based on the assumption that ice grows in the direction normal to the wing surface. The rime ice accretion on a GLC-305 swept wing model under some specific conditions has been simulated to validate the present method. Furthermore, different flight conditions, namely, different angles of attack and different angles of sideslip, have been dealt with to investigate their effects on rime ice accretion as well as the corresponding aerodynamic effects.

Shock-tunnel experiments are carried out to study the strong interaction between an impinging shock wave and boundary layer on a flat plate, accompanied by large separation bubble with a length comparable to the distance of the location of shock impingement from the leading edge of the plate. For nominal freestream Mach numbers ranging from 6 to 8.5, moderate to high total enthalpies of 1.3MJ/kg to 6MJ/kg are simulated in the Indian Institute of Science's hypersonic shock tunnels HST-2 (a conventional Hypersonic Shock Tunnel) and Free Piston Shock Tunnel (FPST) with freestream Reynolds numbers ranging from 4 × 106/m to 0.3 × 106/m. The strong impinging shock is generated by a wedge (or shock generator) at an angle of 30.96° to the freestream. From the time-resolved Schlieren flow visualisations using a high-speed camera and surface pressure measurements on the flat plate using fast response sensors, a statistically steady flow field with a large separation bubble was established within the short test time of the shock tunnels (around 600µs in HST-2 and 300µs in FPST). The role of various parameters on the interaction – Mach number, location of shock impingement and flow total enthalpy – are investigated from the measured separation length and surface pressure distribution. For the nominal Mach number of 8.5, with shock impingement at 100mm from the leading edge, the separation length increased from 60mm to 70mm as the total enthalpy is increased from 1.6MJ/kg to 2.4MJ/kg; whereas it dropped drastically to 30-40mm at 6MJ/kg. This is due to the prominence of real gas effects at higher enthalpies.

Rocket motor nozzle flow geometry is considered through its influence on the thrust vector control (TVC) performances. Extensive research is conducted using theoretical and software simulations and compared with experimental results. Cold and hot flow test equipments are used. The main objective of the research is to establish the methodology of flow geometry optimisation on the TVC hardware system. Several geometry parameters are examined in detail and their effects on the system performances are presented. The discovered effects are used as guidelines in the TVC system design process. A numerical method is presented for the determination of dynamic response time upper limit for the TVC system based on the gas flow dynamics performances.

In this paper, sliding mode control and disturbance observer are used to design a new continuous composite guidance law with terminal angle constraint. The robustness and finite-time convergence of the proposed guidance law is established using the Lyapunov stability theory. For performance improvement, a nonlinear disturbance observer, which can be viewed as a ‘patch’ for the original guidance law, is designed to estimate the target manoeuvre. Theoretical analysis and simulation results demonstrate the effectiveness of the proposed method.