Adverse meteorological conditions often contribute to the formation of ice on aircraft wing section, engine nacelle and other parts leading to the loss of lift coefficient and increase in drag coefficient affecting aircraft control and stability. This paper addresses the problem of in-flight icing on an asymmetric aerofoil under three different ambient and cloud conditions. The study involves prediction of the leading-edge ice thickness using a numerical model developed from the mass and energy conservation law and Messinger freezing fraction model at the same Reynolds number. Later on, degradation in the aerodynamic performance of the iced aerofoil was also investigated using the computational fluid dynamics (CFD) technique, taking the flow field around a 2D aerofoil geometry into account. The aerodynamic study indicates that cumulus clouds embedded with stratified clouds contribute to the formation of mixed ice on aerofoil leading edge and causes the worst icing scenario reducing the lift coefficient to 90% and increasing the drag coefficient to 800% for the same ambient conditions.

With the aviation industry facing increasing environmental and energy challenges, there has been a growing demand for sustainable aviation fuel (SAF). Previous studies have shown the role of ice accretion, release and blockage in aviation-related incidents and accidents with conventional jet fuel. However, there is no available data that establishes the magnitude of influence new fuel compositions will pose on ice formation and accretion in aircraft fuel systems. A recirculating fuel test rig capable of cooling fuel from ambient to −30°C within 4h was built by Airbus to simulate conditions in an aircraft wing tank and allow characterisation of ice accretion. The key characteristic was the pressure drop across an inline fuel strainer for the different SAF explored but visual analysis of ice accretion on the strainer mesh (filters used in protecting fuel feed pumps) was also performed for individual experimental runs for comparison. Measurements revealed that 100% conventional fuel exhibited a higher propensity to strainer blockage compared to the SAF tested. However, all SAF blends behaved differently as the blending ratio with Jet A-1 fuel had an impact on the pressure differential at different temperatures. Data from this work are essential to establish confidence in the safe operation of future aircraft fuel systems that will potentially be compatible with 100 % SAF.

The reduction of computational costs in the context of the Multidisciplinary Design Optimisation of a typical medium-range aircraft was investigated through an assessment of active constraints and the use of multi-fidelity models-based estimation of drag and structural stress. The results show that for this problem, from the set of considered constraints that includes flutter boundary, the active constraint is a 2.5g pull up Maximum Take Off Weight. Results show that the multi-fidelity approach reduced the required high-fidelity aerodynamic number of evaluations, for both drag assessment and stress assessment with sufficient level of accuracy for the former and conservatively for the latter. Further computational cost reduction can be achieved using a surrogate model based Multidisciplinary Design Optimisation. The best configuration attained shows an Aspect Ratio increase of 16%, a reduction of 4.5% in fuel consumption and wing structural weight increase of 2.7% relative to a predefined baseline configuration.

The performance of hypersonic vehicles in the take-off stage considerably influences their capability of accomplishing the flight tasks. This study is aimed at enhancing the take-off performance of a cruise aircraft using the improved chimp optimisation algorithm. The proposed algorithm, which uses the Sobol sequence for initial population generation and a function of the weight factors, can effectively overcome the problems of premature convergence and low accuracy of the original algorithm. In particular, the Sobol sequence aims to obtain a better fitness value in the first iteration, and the weight factor aims to accelerate the convergence speed and avoid the local optimal solution. The take-off mass model of the hypersonic vehicle is constructed considering the flight data obtained using the pseudo-spectral method in the climb phase. Simulations are performed to evaluate the algorithm performance, and the results show that the algorithm can rapidly and stably optimise the benchmark function. Compared to the original algorithm, the proposed algorithm requires 28.89% less optimisation time and yields an optimised take-off mass that is 1.72kg smaller.

Leakage, deformations, power loss, heat generation in the support seal system and other issues are typical when support seals are developed. The design of the support seal system has progressively evolved over recent decades as part of an ongoing effort to provide effective cooling for the aero engine secondary air system. In particular, oil heat management in the oil chamber has strict requirements, which limit the heat generation of the support seal system. The potential of supporting seal research with an oil system is investigated in this work. The combination of the CFD/FEA method and quantifying the heat generation entering the oil chamber allows for improvements not just to the individual buffer air seal unit, but the oil seal together. The analysis relies on the combination of quantifying heat generation entering the oil chamber to provide a mutual influence of neighbouring labyrinth seals. The mutual influence requires further analysis, considering the thermal deformation of the rotor/stator to provide further accurate geometry parameters in preliminary seal designs. The experimental test was conducted to verify the preliminary CFD-FEA loosely coupled analysis result, which reveals that in a turbine support seal system, the radius of the buffer air seal has a significant influence on the leakage flow rate and power loss of the oil seal, which should take into account the integral influence of the pressure difference of the oil seal caused by the radius change of the buffer air seal and the running clearance of the oil seal.

The ability to perform air-to-air refuelling (AAR) can dramatically extend the utility of helicopters, through effectively providing unlimited range. For helicopters, AAR is typically performed utilising the probe-and-drogue aerial refuelling method. This is a complex manoeuver, where normally both the helicopter and tanker aircraft are operating at the limits of their flight envelopes. In addition, the wake flow from the tanker aircraft can cause a significant disturbance on the refuelling helicopter. This paper presents the initial evaluation of an AAR scenario constructed within DLR’s flight simulator, the Air Vehicle Simulator (AVES), based on current procedures and pilot interviews. A mission task was defined to assess the scenario in AVES and results are subsequently discussed. For pilots unfamiliar to formation flight or HAAR, the results show the difficulty of the flying task itself at the given cueing. Measures for improvement in future investigations are suggested.

Helicopter component load estimation can be achieved through a variety of machine learning techniques and algorithms. A range of ensemble integration techniques were investigated in order to leverage multiple machine learning models to estimate main rotor yoke loads from flight state and control system parameters. The techniques included simple averaging, weighted averaging and forward selection. Performance of the models was evaluated using four metrics: root mean squared error, correlation coefficient and the interquartile ranges of these two metrics. When compared, every ensemble outperformed the best individual model. The ensembles using forward selection achieved the best performance. The resulting output is more robust, more highly correlated and achieves lower error values as compared to the top individual models. While individual model outputs can vary significantly, confidence in their results can be greatly increased through the use of a diverse set of models and ensemble techniques.

This paper presents progress towards a transition modelling capability for use in the numerical solution of the Reynolds-averaged Navier-Stokes equations that provides accurate predictions for transonic flows and is thus suitable for use in the design of wings for aircraft flying at transonic speeds. To this end, compressibility corrections are developed and investigated to extend commonly used empirical correlations to transonic flight conditions while retaining their accuracy at low speeds. A compressibility correction for Tollmien-Schlichting instabilities is developed and applied to a smooth local correlation-based transition model and a stationary crossflow instability compressibility correction is included by adding a new crossflow source term function. Two- and three-dimensional transonic transition test cases demonstrate that the Tollmien-Schlichting compressibility correction produces substantially improved agreement with the experimental transition locations, particularly for higher Reynolds number applications where the effects of flow compressibility are expected to be more significant, such as the NASA CRM-NLF wing-body configuration, while the crossflow compressibility correction prevents an inaccurate, upstream transition front. The compressibility corrections and modifications do not significantly affect the numerical behaviour of the model, which provides an efficient alternative to non-local and higher-fidelity approaches, and can be applied to other transport-equation-based transition models with low-speed empirical correlations without affecting their predictive capability in the incompressible regime.

This paper consists of CFD and experimental study for shell projectile at angle of attack 00 and at various Mach numbers (0.7, 0.9, 1.0, 1.5 and 2.0), without spin effect. Passive method of base modification is used to reduce base drag. The goal of this study is to reduce base drag by utilising a base bleed approach called the Inward Turning Base Bleed Method (IWTB). The concept of IWTB is to draw relatively high-pressure air behind the driving band and allow it to pass through the bleed holes and direct it into the low-pressure base. This will in turn raise the base pressure and lower the base drag. The study comprises of (i) basic model with a boattail angle of $8^{\circ}$ , (ii) eight different cavity models having different lip thickness and depth thickness ratios and (iii) nine different IWTB configurations for the optimised cavity model. At first, cavity model is optimised by varying the lip thickness and depth thickness ratio. Later IWTB parameters such as the bleed hole entry angles $10^{\circ}$ , $12^{\circ}$ & $15^{\circ}$ , exit bleed hole angles $30^{\circ}$ , $45^{\circ}$ & $60^{\circ}$ , bleed hole diameters 3mm, 4mm, 5mm & 6mm and number of bleed holes i.e., 4 & 8 were studied and optimised. Based on this study, the model with 8 holes, 3mm base bleed hole diameter, $15^{\circ}$ entry and $60^{\circ}$ exit angle gives comparatively lower base drag. CFD result shows, in supersonic region the base drag reduction for optimised shell projectile is $3.08\%$ ; at Mach number 0.9 the base reduction is $75.63\%$ ; and in subsonic region it reduces to $23.53\%$ . CFD results were compared with experimental result, is found to be good and the differences lies within $3.85\%$ .

The aim of this paper is to present a plane-stress damage model based on the Classical Lamination Theory (CLT), developed for polymer fibre-based composite. The proposed numerical model utilises a damage mechanics methodology coupled with fracture mechanics to predict composite failure, particularly under quasi-static and dynamic loadings. In addition, the proposed constitutive equations consider a single secant modulus to describe its tensile and compressive modulus, as opposed to the physically proposed tier models for polymer fibres which possesses a ‘skin-core’ structure. The result of single element and coupon-level modelling showed excellent correlation with the experimental results. In addition, it was also found that the proposed numerical model showed considerable accuracy on the response of the composite under low and high velocity impact loadings.

In some small airports, a parallel taxiway is not built due to space restrictions or financial issues; hence, the runway itself is often used as a taxiway in this type of airport. After touch down, aircraft move to the U-turn area at the end of the runway and turn 180 degrees, then move back to the desired point, such as a gate or the apron, using the runway. The runway is blocked to other arrivals and departures during this process. This process, called backtrack or back-taxi, can result in high delays for both arrivals and departures. Runway occupancy times (ROTs) vary depending on numerous conditions, including pilot performance, weather conditions, aircraft type, etc. Although there are speed restrictions and procedures announced in advance, the actual performance can be uncertain. In addition, most aircraft can make a U-turn as soon as they sufficiently reduce their speed before they reach the U-turn area especially if they are already delayed. These situations bring enormous uncertainties for traffic management at such an airport. Controllers may need help to sequence aircraft, particularly in busy traffic. In this study, a stochastic mathematical model is developed to sequence arrival/departure operations at such an airport considering the ROT uncertainties of arrivals. The objective function of the developed model is determined as the minimisation of the total delay. ROT data was obtained by observing radar tracks of 120 arriving flights. Reasonable ROT scenarios with various probabilities to represent ROT uncertainties were integrated into the mathematical modeling. In addition, two different sequencing approaches are presented as well as the first come first serve (FCFS) approach. As a result, the proposed stochastic approach provides robust sequences applicable for all ROT scenarios with significant delay savings up to an average of 18.4% and 39.5% compared to deterministic and FCFS approaches, respectively.

Hybrid-electric propulsion is a promising alternative to sustainable aviation and is mainly considered for the commuter and regional aircraft class. However, the development of hybrid-electric propulsion variants is affected by the technology readiness level of electric components. The components’ technology will determine the electrification benefit, compared to a conventional aircraft, and will suggest which is the most beneficial variant and which has a closer entry into service date. Within this work, three different dates are explored, namely 2027, 2030 and 2040, to size three Parallel and three Series hybrid-electric architecture variants using an in-house aircraft sizing tool. All variants are compared to a conventional configuration sized using technological assumptions of 2014, with the main comparison metrics being the aircraft block fuel, energy consumption, direct operating cost and holistic environmental impact. On one hand, the Parallel configurations have reduced maximum take-off mass and mission energy consumption compared to the Series, however, the latter show a greater potential for block fuel reduction and require less onboard energy for the same mission. The annual operating cost evaluation indicates that the Parallel hybrid variant of 2030 has greater operational costs than the respective Series variant; however, it has reduced capital costs compared to the latter, making it more economical to operate considering both costs. Additionally, in the case of an energy recession, both hybrid variants of 2030 show a further cost reduction, with the Series having a total reduction of 10.4% excluding capital costs, compared to the reference aircraft. Moreover, the life cycle assessment shows that the Series variants have a lower environmental impact, both compared to the reference aircraft and the Parallel variants. The former could be up to 59.7% less detrimental to the environment than the reference aircraft, whereas the latter up to 23.9%, with the integration of renewable sources for electricity production. Finally, by the year 2040, the Series variant shows outstanding performance in all comparison metrics, compared to the Parallel and the reference aircraft.

This paper presents the results of the experimental study carried out to address the issues of base heating and smooth separation of the stage of launch vehicles. The pressure at the base of a convergent-divergent circular nozzle, from which Mach 1.8 jet emanates, attached to an annular shroud of larger area is controlled by providing air vents on the shroud. On the shroud, vent holes were made at different azimuthal locations, to entrain the surrounding air mass at a higher pressure, pa, to increase the low-pressure, pb, at the base region, caused by the suction creating large-scale vortices formed owing to the sudden expansion of the jet emerging from the nozzle into the shroud. For different number and size of the vents on the shroud, the base pressure was measured. This measurement was done at five levels of overexpansion of the nozzle in the range from –64% to –58%. It is found that increase in vent area results in increase of base pressure, up to some limiting level of the area. Also, the increase of base pressure for the case of vents closer to the nozzle exit is found to be marginally more than the increase caused by vents at distances away from the nozzle exit. Increase of base pressure can be regarded as an advantage not only from base heating point of view but also from the point of view of deflection of the plume to the shroud wall for uniform melting of the pyro layer bonding the stages of the launch vehicle, leading to a smooth separation of the launch vehicle stages.

Replacing kerosene with hydrogen as a fuel for propelling aeroplanes will undoubtably have ramifications for the existing certification basis and the necessary means of compliance. What this paper attempts to do is explain how a future hydrogen-powered aeroplane could be certified and how this will influence the aircraft’s design. Emphasis has been placed on the safety of hydrogen with regards flammability, ignition energy, cryogenic states and crashworthiness. The influence of certification on a future hydrogen-fuelled aeroplane’s configuration has been established. Lessons learnt from rockets are limited in scope. Consideration of hydrogen’s cryogenic and ambient temperature properties has established some fundamental safety principles. Fuel system crashworthiness influence on certification has been examined. An explosion prevention certification scheme has been developed. It is proposed to exclude oxygen from the fuel system and avoid accumulation of leaked or permeated hydrogen from adjacent zones. Three generic hydrogen-fuelled configurations have been considered for certification: adapting conventional airliner design, hydrogen-electric (fuel cell) propulsion and an unconventional approach. Hydrogen as a safe gas turbine fuel has been examined in technical and environmental terms. Further experimental work to close knowledge gaps in the aviation use of hydrogen has been suggested. In conclusion the necessary changes to the certification basis are outlined. Comments as to what the configuration of a future hydrogen-fuelled aeroplane might look like are provided. These observations are based on analysis of hydrogen’s properties, the existing certification basis, knowledge gaps, and what Special Conditions and Acceptable Means of Compliance might emerge in a future design Type Certification.

Flutter suppression is an important measure to improve fatigue life and enhance the performance of aircraft in modern aircraft design. In order to design more effective controllers for flutter suppression with high efficiency, an efficient reduced-order framework for active/passive hybrid flutter suppression is proposed. The traditional CFD-based ROMs have been successfully applied to active flutter suppression with high accuracy and efficiency. But, when a structure modification is made such as in aeroelastic tailoring and aeroelastic structural optimisation, the structural model should be updated, and the expensive, time-consuming CFD-based ROMs have to be reconstructed; such a process is impractical for passive flutter suppression. To overcome the realistic challenge, an efficient reduced-order framework for active/passive hybrid flutter suppression is proposed by extending an efficient aeroelastic CFD-based POD/ROM which we have developed. The proposed framework is demonstrated and evaluated using an improved AGARD 445.6 wing model. The results show that the proposed framework can accurately predict the aeroelastic response for active/passive hybrid flutter suppression with high efficiency. It provides a powerful tool for active/passive hybrid flutter suppression, and therefore, is ideally suited to design more effective controllers, and may have the potential to reduce the overall cost of aircraft design.

This paper considers the problem of a three-axis flexible satellite attitude stabilisation subject to the vibration of flexible appendages and external environmental disturbances, which affect the rigid body motion. To solve this problem, a disturbance observer is proposed to estimate and thereby reject the flexible appendage vibration. Based on the H∞ and Linear Matrix Inequality (LMI) approach, a controller for spacecraft with flexible appendages is proposed to ensure robustness as well as attitude stability with high precision. Stability analysis of the overall closed-loop system is provided via the Lyapunov method. The simulation results of three-axis flexible spacecraft demonstrate the robustness and effectiveness of the proposed method.

The need for innovative solutions to enable aerial platforms to fly faster, higher, and longer continues to remain a primary focus for airframe designers. This paper outlines work undertaken to apply a morphing wing warping technology onto a generic Unmanned Aerial Vehicle to deliver enhanced flight performance, efficiency and control capabilities. The prototype employs wings of novel construction which provide both near resistance-free compliance in twist as well as adequate structural stiffness to resist applied loads; all while preserving an aerodynamically smooth surface. Used in combination with developed and integrated closed-loop feedback control architecture, a real-time, non-linear, span-wise wing twist adjustment capability required for optimised flight under differing operating conditions and flight requirements, is demonstrated. Experimental results obtained from a wind tunnel test program show up to a 72% increase in lift to drag ratio under certain conditions compared to a fixed baseline providing some confidence that the combination could be used to realise a step change in flight performance.

This paper describes the methodology used to analyse oscillations of foils of a wide range of aspect ratios, 0.5 ≤ AR ≤ 4, and Reynolds numbers, 104 ≤ Re ≤ 105, for energy harvesting purposes. The foils were fixed at their trailing edge, and their dynamical behaviour was captured as the wind speed was varied. The foil response was then analysed as a function of velocity, Reynolds number, oscillation amplitude and frequency. Additionally, the forces and moments acting on the foils were measured, utilising an aerodynamic scale, designed and built in-house. An empirical power generation equation was derived to determine the foil characteristics for maximum energy harvesting production. The results show that a flexible foil with AR = 3 with oscillations in the large-amplitude regime is the most effective for energy harvesting.

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 accuracy of several numerical schemes for solving the lifting-line equation is investigated. Circulation is represented on discrete elements using polynomials of varying degree, and a novel scheme is introduced based on a discontinuous representation that permits arbitrary polynomial degrees to be used. Satisfying the Helmholtz theorems at inter-element boundaries penalises the discontinuities in the circulation distribution, which helps ensure the solution converges towards the correct, continuous behaviour as the number of elements increases. It is found that the singular vorticity at the wing tips drives the leading-order error of the solution. With constant panel widths, numerical schemes exhibit suboptimal accuracy irrespective of the basis degree; however, driving the width of the tip panel to zero at a rate faster than the domain average enables improved accuracy to be recovered for the quadratic-strength elements. In all cases considered, higher-order circulation elements exhibit higher accuracy than their lower-order counterparts for the same total degrees of freedom in the solution. It is also found that the discontinuous quadratic elements are more accurate than their continuous counterparts while also being more flexible for geometric representation.