There are over 16,000 pieces of debris that are larger than 10cm in diameter being tracked which could cause a catastrophic destruction of a satellite if they were to collide. To reduce the likelihood of this happening, debris production is being reduced through debris mitigation techniques; however the debris already in space will need to be removed as well. To design an effective debris removal mission, the first logical step would be to gather as much detail as possible on debris that is known to exist within the orbital band of interest. This report presents a catalogue of debris, created entirely from open source data, for an orbital band within the 900-1,000km altitude range and 82-83° inclination, and offers a possible candidate for deorbit – the Cosmos 3M final-stage rocket body, of which there are 142, each with a mass of 1,435kg, length of 6·5m and diameter of 2·4m.

This paper presents a modified one dimensional model for the pulsed plasma thruster. The model takes into account the impact of thruster geometry and capacitor initial energy on plasma inductance and resistance, as well as the ablated mass per pulse. These three factors have significant impact on the performance of the thruster and have not been considered concurrently before. The model is verified by comparing the estimated specific impulse, impulse bit and thrust-to-power ratio with experimental measurements for space qualified thrusters. The maximum error is estimated to be about 8% when using the modified model, while the conventional snowplow model produces error up to about 70%. The modified model is then employed to generate design charts that would help in selecting the proper combination of thruster geometry and energy level in the preliminary design phase of a pulsed plasma thruster for different space missions.

The paper describes a method of representing damage on a wing due to multiple warhead fragments, and investigates two of the key variables: fragment impact density and hole diameter. The aerodynamic effects of the damage were quantified by wind-tunnel tests on a two-dimensional wing at a Reynolds number of 5 × 105. The wing was of hollow construction with leading and trailing-edge spars. In all of the cases tested, simulated fragment damage resulted in significant lift losses, drag increases and pitching moment changes. Increasing fragment density or hole size resulted in greater effects. To a first order approximation, both lift and drag increments at a given incidence were related to the percentage wing area removed. Surface flow visualisation showed that low fragment densities and small damage sizes resulted in a complex flow structure on the surface of the wing. This was made up of boundary-layer growth between the damage holes, attached wakes from the forward damage holes and separated surface flow over the rear of the wing. For these cases, individual hole patterns showed similar flow mechanisms to those seen for larger scale gunfire damage cases. Increased fragment density and hole size resulted in upper surface flow separation at the first row of holes. Behind this separation, the flow was attached and consisted of the combined wakes from the forward damage holes. Investigations into the influence of internal model structure indicated that trends in coefficient changes were similar for both hollow and solid wings. However, the magnitudes of the effects were found to be smaller for hollow wings than for solid wings.

Computational simulations have been used to study a new low-boom, axisymmetric, external compression supersonic inlet with an on-design Mach number of 1·7. The inlet incorporates a relaxed compression surface with a near-zero cowl angle to help reduce external oblique shock waves due to spillage and cowl geometry. To reduce mechanical complexity the inlet is designed with zero-bleed. To understand the impact on performance and shock overpressure caused by the inlet itself, several throat and diffuser designs were simulated. The computations utilised a Reynolds-averaged Navier-Stokes code. Inflow properties were held consistent with the operational characteristics of the NASA GRC 8′ × 6′ Supersonic Wind Tunnel (SWT) for experimental testing of the inlet of Mach 1·67 at a Reynolds number of 5·4 × 106. Stagnation pressure recovery performance for the baseline condition exceeded 94% at design mass flow rates, and reduced only slightly with increases in Mach number (consistent with theoretical predictions) and extension of cowl position. The simulations also showed that the relaxed compression surface combined with the near-zero cowl angle helps to significantly reduce external oblique shocks This is partly due to the reduced inlet spillage in combination with a reduced overall turning angle placed on the free-stream flow relative to the cowl shape.

This work presents a study of acoustic oscillations generated and the wave structure associated with supersonic flow past wall mounted 3D open cavities of varying length-to-width (L/W) ratio. Experiments were conducted to investigate the acoustic signature generated by the cavities at freestream Mach number of 1·7. The effect of L/W ratio of the cavity on the dominant modes of the acoustic signature registered on different walls of the cavities is investigated for an L/W range of 0·83-4. Shift in the dominant acoustic mode is observed as L/W ratio changes from 3 to 4. Statistical analysis of pressure data showed existence of acoustic waves and spreading of acoustic energy over different modes with change in cavity width. Time averaged schlieren visualisation indicated variation of shock and shear layer structure in the mainstream for the different cavities. Acoustic waves generated by the presence of the cavity and the dynamic behaviour of the shear layer were observed during instantaneous shadowgraph visualisation. Numerical simulation was done to make a prior assessment of the flow structure and the results are in good agreement with those from experiments. Ratio of mass exchange between cavity and mainflow and the cavity volume was observed to have profound effect on the magnitude of pressure oscillations generated by the cavities.

Modern gas turbine combustor design is a complex task which includes both experimental and empirical knowledge. Numerous parameters have to be considered for combustor designs which include combustor size, combustion efficiency, emissions and so on. Several empirical correlations and experienced approaches have been developed and summarised in literature for designing conventional combustors. A large number of advanced technologies have been successfully employed to reduce emissions significantly in the last few decades. There is no literature in the public domain for providing detailed design methodologies of triple annular combustors.

The objective of this study is to provide a detailed method designing a triple annular dry low emission industrial combustor and evaluate its performance, based on the operating conditions of an industrial engine. The design methodology employs semi-empirical and empirical models for designing different components of gas turbine combustors. Meanwhile, advanced DLE methods such as lean fuel combustion, premixed methods, staged combustion, triple annular, multi-passage diffusers, machined cooling rings, DACRS and heat shields are employed to cut down emissions. The design process is shown step by step for design and performance evaluation of the combustor.

The performance of this combustor is predicted, it shows that NOx emissions could be reduced by 60%-90% as compared with conventional single annular combustors.

This paper is concerned with the development of a failure initiation and progressive failure analysis (PFA) method for advanced composite structures. The present PFA model is capable of predicting interactive out-of-plane and in-plane failure modes observed in fiber reinforced composite laminates including interlaminar behavior and matrix microdamage at the mesoscale. A probability analysis tool is coupled with the PFA to account for uncertainty in modelling parameters caused by material variability and manufacturing inconsistencies. The progressive damage response of a laminated composite panel with an initial delamination is studied and used to demonstrate the PFA modelling framework that is presented here.

This work considers the effects of camber morphing, both in magnitude and chord position, on the performance of a generic unmanned air vehicle (UAV). The focus is to maximise appropriate aerodynamic factors across the mission by optimising the wing camber. Specifically, the enhancement of range, endurance, and stall speed is sought by means of maximising their aerodynamic performance parameters, CL/CD, CL3/2/CD, and CLmax respectively. An analysis of the effects of camber morphing is carried out using the vortex panel code, XFOIL, utilising aerofoils from the NACA four-digit family. The results are then adjusted to account for 3D flow factors such as induced drag, offering a more realistic appraisal of their effectiveness. Flight testing is then performed on four wings of fixed aerofoil sections, optimised for each performance characteristic, to validate the trends observed in the XFOIL data onboard a 1·64m span aircraft.

The purpose of this paper is developing an efficient flight control strategy in terms of time response characteristics, robustness with respect to both parametric uncertainties and un-modeled nonlinear terms, number of required measurements, and computational burden. The proposed method is based on combination of a classic controller as principal section of the autopilot and a multi-objective genetic algorithm-based fuzzy output sliding mode control (FOSMC). FOSMC not only modifies robustness of the classic controller against uncertainties and external disturbances, but also modifies its time response for wide range of commands. FOSMC is a single input-single output controller that is based on the system output instead of the system states. In this situation, the proposed autopilot does not require measurement of other variables and observer, and also it is practicable because of considerable reduction in rule inferences then computational burden. As a critical application, the proposed method is applied to design the altitude hold mode autopilot for an UAV which is non-minimum phase, uncertain, and nonlinear.

The flow past an unmanned combat air vehicle (UCAV) 1303 was investigated in a water tunnel in an effort to understand the complex flow field. The UCAV 1303 is characterised by a non slender, 47° swept-leading-edge delta wing, with a cranked trailing edge, a cropped wing tip and a fuselage. It also has no vertical tail. Dye visualisation pictures and five-axis strain gage load data were obtained on a 1:72 scale model in both steady flow and while the model executed preprogrammed manoeuvres. From the strain gage data, body axis force and moment data were derived. In addition to some similarities, comparison of the flow features observed with those of more well known slender delta wing flows showed some remarkable differences such as tip-stall, the presence of two like-sense vortices on the same side of the wing, their simultaneous spiral and bubble bursting. This paper predominantly discusses the flow visualisation results and attempts to correlate the load data to some of the observed flow features.

Prediction of the buoyancy of a thermal balloon, or Montgolfière, intended to float in the atmosphere of Titan is discussed. A laboratory based experiment designed to measure the buoyant lift of an electrically-heated, single-wall, natural-shape balloon is described. The experimental results presented closely match an analytical model employing established heat transfer correlations with selected parameter values. When the model is extrapolated to the cryogenic conditions representative of Titan’s lower troposphere, using the same correlations and parameters, the estimated buoyancy is found to be substantially higher than has been previously predicted. To account for the buoyancy difference, it is suggested that the internal free convection heat transfer rate is significantly lower than has been assumed in previous studies. To substantiate this result, it is recommended that further experiments should be performed at higher Rayleigh numbers.

In this paper, probabilistic considerations are introduced in a model for fatigue life prediction for a riveted lap joint using Paris’ law. Initial crack sizes are distributed according to a truncated lognormal distribution, which is chosen to avoid known complications due to Paris’ law. The stress intensity factor for a single rivet hole is calculated, and is generalized to a lap joint. The probability of the existence of a crack in two domains of interest are evaluated, and the effect of a single inspection, modeled using the Probability of Detection, is studied. Additionally, the probability of detection concept is extended by linking it to applied stress and number of elapsed cycles using Bayes’ theorem and ramifications are explored.

This paper presents the experimental results of the role of struts in supersonic mixing. Experiments were carried out with novel strut configurations to show their capabilities on mixing with reasonable total pressure losses. The performances were compared with the Baseline Strut configurations (BSPI and BSNI). The analysis presented includes the mixing quantifications using Mie scattering signature, flow field visualisation, measurement of wall static pressure and the total pressure loss calculations. The results clearly demonstrated that the proposed strut configurations achieved increased mixing (7-8%) compared to BSPI with increase in total pressure loss (2%). On the other hand, when compared with BSNI, the mixing performance was found to be decreased by 6% with reduced total pressure loss (12%).

An experimental investigation is presented to validate endurance estimates for a battery powered flight vehicle. In this study a radio controlled model aircraft was ‘flown’ inside a wind tunnel. Various Lithium Ion batteries were used to power the model. Battery electrical characteristics were recorded during each run to assess battery discharge behaviour. Theoretical estimates based upon battery characteristics yielded endurance estimates within 3% of that measured experimentally. The discharge characteristics of the batteries were found to be the largest source of uncertainty for endurance estimates.

New aircraft require years of development from concept to realisation and can be prone to delays. Consequently, military operators take existing fleets and operate them in a different role. The objective of this study is to examine the effect of operating a typical low bypass military fast jet engine, originally designed for a European theatre, in a hot and harsh climate. The specific purpose is to determine the effect on the high-pressure turbine blade life and the life- cycle cost of the engine.

A mission profile and respective performance conditions were analysed and modelled using an in-house performance tool. The flow conditions were simulated using ANSYS® FLUENT. A conjugated heat transfer solution was adopted to determine the blade metal temperature. The blade was modelled physically in 3D using SIMULIA® ABAQUS FEA software. The stresses were derived and used to calculate the temperature coupled low cycle fatigue and creep life.

A deterioration case was also studied to evaluate the effect of sand and dust ingestion. There was a significant life reduction of approximately 50% due to creep. The reduction in life was inversely proportional to the life cycle cost of the engine depending on the operating conditions. The results were compared with similar engines and summarised in the context of airworthiness regulations and component integrity.

This paper presents an overview of an advanced, conceptual wing-box weight estimation and sizing model for transport aircraft. The model is based on linear thin-walled beam theory, where the wing-box is modelled as a simple, swept tapered multi-element beam. It consists of three coupled modules, namely sizing, aeroelastic analysis, and weight prediction. The sizing module performs generic wing-box sizing using a multi-element strategy. Three design cases are considered for each wing-box element. The aeroelastic analysis module accounts for static aeroelastic requirements and estimates their impact on the wing-box sizing. The weight prediction module estimates the wing-box weight based on the sizing process, including static aeroelastic requirements. The breakdown of the models into modules increases its flexibility for future enhancements to cover complex wing geometries and advanced aerospace materials. The model has been validated using five different transport aircraft. It has shown to be sufficiently robust, yielding an error bandwidth of ±3%, an average error estimate of -0·2%, and a standard error estimate of 1·5%.

The following paper presents detailed aerodynamic data of a Scottish Aviation Bulldog light aircraft. The data is taken from the pre-stall region of the aircraft flight envelope through two flight test methods and from a geometrically accurate computational fluid dynamics (CFD) model of the full scale aircraft, which was meshed in Ansys ICEM CFD and solved in Ansys Fluent. The fidelity of the CFD model was achieved by development of a CATIA solid model with surfaces matching a spatial point cloud of the aircraft taken using a 3D laser scanner. Following a CFD verification process, a 3·4m hybrid mesh with a Spalart-Allmaras (SA) turbulence model was found to give the best overall lift and drag characteristics. Further detailed comparisons with the glide flight test data showed the CFD drag polar to have 63% lower zero lift drag, although this discrepancy was related to the simplification of the original CATIA surface model, which excluded the undercarriage, aerials and other protuberance drags. Inclusion of estimates of these sources of drag resulted in a match in zero lift drag to within 15% and a maximum lift to drag of 10:1 which was within 11% of the glide flight test result. The remaining drag discrepancy is attributed to other effects including trim drag and the surface finish of the actual aircraft.

A procedure for defining an empirical formula for the mass estimation of the fore and aft wings field Uof a medium range box wing aircraft is described. The procedure is based upon the work of Howe for estimating the wing mass of conventional cantilever wing aircraft. The paper outlines the procedure used to relate conventional cantilever wings to box wing aircraft wings. Using a vortex lattice tool, finite element methods and regression analysis, the modification performed on the coefficient in Howe’s method to enable its use on a medium range box wing aircraft is outlined. The results show that the fore and aft wings would use the same correction coefficient and that the aft wing would therefore be lighter than the fore wing on a medium range box wing aircraft.

A computational study has been carried out to analyse the shock wave turbulent boundary layer interaction in a two-dimensional compression ramp flow for a free stream Mach number of 2·94. Ramp angles ranging from 8° to 24° are used to produce the full range of possible flow fields, including flows with minor separation, moderate separation, and significant amount of separation. The model has been analysed using 2D numerical simulations based on a commercially available Computational Fluid Dynamics (CFD) Code, Fluent, that employs k-ω Shear Stress Transport (SST) turbulence model. The computed data for surface pressure distribution indicated a good agreement with experiment. Numerical results obtained through the present series of computations indicate an increased extent of separated zone, and thus show increased upstream influence when compared to experiment. Total pressure loss has shown to increase downstream of separation location and increase when corner angle increases. Secondary separation has been observed for higher angles.