Thermal-mechanical design is a time-consuming process even at its preliminary design stage. This is due to the large number of components and boundary condition data, the complexity of the geometry, and the iterative nature of the design process. The conventional design process separates the geometric and physical models and results in considerable human interventions during the design process. By assigning the breakpoints to engine assembly features as internal parameters, this article reports a novel feature-based design approach where the associated boundary conditions are represented parametrically along the feature geometric contours. They are updated automatically as per the geometrical changes, including topological changes, and hence bridging the gap between the geometric and physical models. The current approach enables data re-use of both the geometries and physical information from previous engine designs to generate new designs, dispensing with the excessive human interventions. Although the methodology is generic and applicable to other design scenarios, its capability is demonstrated in this article by some representative challenging industrial applications, sitting in the 2D preliminary gas-turbine design domain. The test cases show that the method can significantly reduce the time-cost of the iterative thermal-mechanical design flow.

This paper discusses how the processes of Continuing Airworthiness Management Organisations can be made more efficient by applying Lean principles and digital solutions, to obtain costs savings and improved compliance standards. Lean management is the process of eliminating waste by maximising value through balancing workloads and cutting out inefficient processes. Though Lean was initially applied to manufacturing production lines, an opportunity exists to apply the principles of Lean in airworthiness management. One such aspect of airworthiness management is the processing of records which is an essential part of safety and compliance. Software solutions designed with Lean philosophy and focus on airworthiness management processes can provide benefits in managing the flow of records and extracting actionable information from them for cost reduction and improving safety. Further value can be extracted by applying additional software solutions that will lay a strong foundation for machine learning and predictive analytics – taking aircraft maintenance from a reactive model towards a proactive model.

Non-axisymmetric endwalls in turbine stages have shown to be a robust method to improve the performance of turbines in both power generation and aero-derivative applications. Non-axisymmetric endwalls target the control of secondary flows and are designed using detailed computational fluid dynamics coupled with a variety of optimisation algorithms and utilising a number of objective functions according to the engine company or researcher's preference. These numerical predictions are often backed up by detailed measurements in linear and annular cascades and later proven in full-scale engine tests. Relatively little literature is available describing their performance in rotating test rigs or at conditions other than design, apart from that of the authors. This study comprehensively revisits the low-speed, model turbines used in the earlier study, replacing all of the 5-hole probe data with more accurate results and additional hot-film measurements. These results together with computational fluid dynamics solutions are used to show the success of the method across a large incidence range and to compare to earlier cascade results for a similar endwall and blade profile to establish the usefulness of cascade testing in this application. In addition, a comparison to two other off-design studies is made. Results indicate that the endwalls successfully improve the rotor total isentropic efficiency at all test conditions and that the improvement increases with increased turning in the blade row, from 0.5% to 1.8% across the incidence range. The results also compare well to the estimation of isentropic efficiency improvement that can be drawn from the cascade testing which stands at 1.55%.

This paper investigates the synergies and trade-offs between passive aeroelastic tailoring and adaptive aeroelastic deformation of a transport composite wing for fuel burn minimisation. This goal is achieved by optimising thickness and stiffness distributions of constitutive laminates, jig-twist shape and distributed control surface deflections through different segments of a nominal “cruise-climb” mission. Enhanced aerostructural efficiency is sought both passively and adaptively as a means of aerodynamic load redistribution, which, in turn, is used for manoeuvre load relief and minimum drag dissipation. Passive shape adaptation is obtained by embedding shear-extension and bend-twist couplings in the laminated wing skins. Adaptive camber changes are provided via full-span trailing-edge flaps. Optimised design solutions are found using a bi-level approach that integrates gradient-based and particle swarm optimisations in order to tailor structural properties at rib-bay level and retrieve blended stacking sequences. Performance benefits from the combination of passive aeroelastic tailoring with adaptive control devices are benchmarked in terms of fuel burn and a payload-range efficiency. It is shown that the aeroservoelastically tailored composite design allows for significant weight and fuel burn improvements when compared to a similar all-metallic wing. Additionally, the trailing-edge flap augmentation can extend the aircraft performance envelope and improve the overall cruise span efficiency to nearly optimal lift distributions.

Machining-induced surface integrity has an important effect on reliability and service life of the components used in the aerospace industry where titanium alloy Ti-6Al-4V is widely applied. Characterisation of machining-induced surface integrity and revealing its effect on fatigue life are conducive to structural fatigue life optimisation design. In the present study, surface topography, residual stress, microstructure and micro-hardness were first characterised in peripheral milling of titanium alloy Ti-6Al-4V. Then, low-cycle fatigue performances of machined specimens were investigated on the basis of the tension-tension tests. Finally, the effects of surface integrity factors (stress concentration factor, residual stress and micro-hardness) on fatigue performances were discussed. Results show that stress concentration can reduce the fatigue life while increasing the residual compressive stress, and micro-hardness is beneficial to prolonging the fatigue life, but when the surface material of the specimen is subjected to plastic deformation due to yield, the residual stress on the surface is relaxed, and the effect on the fatigue performance is disappeared. Under the condition of residual stress relaxation, the stress concentration factor is the main factor to determine the low-cycle fatigue life of titanium alloy Ti-6Al-4V. While for the specimens with no residual stress relaxation, micro-hardness was the key factor to affect the fatigue life, followed by residual stress and stress concentration factor, respectively.

In conversion flight, the aeroelastic modelling of tiltrotor aircraft needs to consider the unsteady effect of the rotor wake bending due to the rotor tilting. In this paper, the unsteady models of the rotor wake bending and dynamic inflow have been introduced into the aeroelastic modelling of the tiltrotor aircraft in conversion flight by using Hamilton’s generalized principle. The method for solving the aeroelastic stability of tiltrotor aircraft in conversion flight has been established by using the small perturbation theory and the Floquet theory. The influences of unsteady dynamic inflow on trim control inputs and aeroelastic stability of a tiltrotor aircraft in conversion flight were calculated and analysed. The calculation results show that the required collective pitch increases with the pylon tilting forward and the unsteady inflow is trimmed primarily by the lateral cyclic pitch of the rotor. The wake bending unsteady dynamic inflow can obviously reduce the stability of the flapping modes of the rotor, and have no obvious influence on the lag modes of the rotor and the motion modes of the wing. The instability of tiltrotor occurs in the chordwise bending mode of the wing when the pylon tilts to a certain angle in high speed forward flight.

An electromechanical coupling model is established for the space-tethered combination (STC) under microgravity environment after target capture by the tethered robot system (TRS). A linearized dynamic model of the STC is put forward with its controllability and observability as a control system analyzed. A double closed-loop tension control strategy is proposed to mitigate the impact and suing longitudinal vibration caused by the velocity difference between the platform and target. Experiment setup is built on a ground-based flotation platform to investigate the impact of the STC. Results of simulation and experimental validation show that the proposed tension control strategy is responsive and rapid in tension tracking and effectively prevent impact.

An investigation of the flow around an obstacle positioned within the wake of a rotor is described. A flow visualisation survey was performed using a smoke wand and particle image velocimetry, and surface pressure measurements on the obstacle were taken. The flow patterns were strongly dependent upon the rotor height above the ground and obstacle, and the relative position of the obstacle and rotor axis. High positive and suction pressures were measured on the obstacle surfaces, and these were unsteady in response to the passage of the vortex driven rotor wake over the surfaces. Integrated surface forces are of the order of the rotor thrust, and unsteady pressure information shows local unsteady loading of the same order as the mean loading. Rotor blade-tip vortex trajectories are responsible for the generation of these forces.

Employing an analytical method, non-linear low-velocity impact response of carbon nanotube (CNT)-reinforced sandwich cylindrical panels in thermal environments is analysed. Two types of core (i.e. homogenous and functionally graded) are considered for sandwich panels. The face sheets of sandwich panels are multi-layer which consist of CNT-reinforced composite (CNTRC) and metal layers. Micromechanical models are used to estimate the material properties of CNTRCs. A higher-order shear deformation theory with a von Kármán-type of kinematic non-linearity provides the equations of motion. Temperature-dependent material properties are used to include the thermal effects. The equations of motion are solved using a two-step perturbation technique. Existing numerical results in the literature are used to validate the present method. The effect of nanotube volume fraction, material property gradient, impactor initial velocity, geometrical parameters of cylindrical panel, temperature change and edge boundary condition on the impact response of cylindrical panel structures is discussed. The quantitative results and analytical formulations can be helpful in better designing of CNTRC structures subjected to low-velocity impact in thermal environments.

Mach 2 jet from a convergent-divergent elliptical nozzle, of aspect ratio 2, has been controlled with limiting flat and arc tabs. The mixing promoting capability of the flat and arc tabs were studied in the presence of different levels of pressure gradient, at the nozzle exit, corresponding to nozzle pressure ratios of 4 to 8, in steps of one. The geometrical blockage of both the tabs is 5% of nozzle exit area. For the flat tab along the minor axis, the waves in the core become weaker and the core length becomes shorter than the uncontrolled jet, at all the NPRs studied. But the flat tab along the major axis promotes mixing only for some NPRs and retards the mixing for the rest of the NPRs studied. At NPR 5, the flat tab along the minor axis causes the largest core length reduction of 86%. For circular arc tab, along the minor axis, the maximum core length reduction is 55% at NPR 6. Arc tab along the major axis protects the core length for the entire range of the nozzle pressure ratios tested and maximum extension in core length is found at NPR 4, which is 40%.

Over the past decade, NASA, under a succession of rotary-wing programs, has been moving towards coupling multiple discipline analyses to evaluate rotorcraft conceptual designs. Handling qualities is one of the component analyses to be included in such a future Multidisciplinary Analysis and Optimization framework for conceptual design of Vertical Take-Off and Landing (VTOL) aircraft. Similarly, the future vision for the capability of the Concept Design and Assessment Technology Area of the U.S Army Aviation Development Directorate also includes a handling qualities component. SIMPLI-FLYD is a tool jointly developed by NASA and the U.S. Army to perform modelling and analysis for the assessment of the handling qualities of rotorcraft conceptual designs. Illustrative scenarios of a tiltrotor in forward flight and a single-main rotor helicopter at hover are analysed using a combined process of SIMPLI-FLYD integrated with the conceptual design sizing tool NDARC. The effects of variations of input parameters such as horizontal tail and tail rotor geometry were evaluated in the form of margins to fixed- and rotary-wing handling qualities metrics and the computed vehicle empty weight. The handling qualities Design Margins are shown to vary across the flight envelope due to both changing flight dynamics and control characteristics and changing handling qualities specification requirements. The current SIMPLI-FLYD capability, lessons learned from its use and future developments are discussed.

To investigate the control effect of the synthetic jet on the aerodynamic characteristic of rotors, a numerical simulation procedure for the rotor flowfield is established. First, a moving-embedded grid method and an unsteady Reynolds Averaged Navier–Stokes (URANS) solver are established for predicting the complex flowfield of rotors. A velocity jet boundary condition over the jet actuator orifice is constructed, and a numerical method for simulating the active flow control on rotors is developed. Then, the effectiveness of the simulation method is validated by comparing the numerical results of jet control on NACA 0015 aerofoil with the experimental data. At last, the aerodynamic characteristic of rotors with synthetic jet actuators located on the suction surface of the blade in forward flight is calculated. The results indicate that the synthetic jet has the capability of improving the aerodynamic characteristic of rotors, especially in inhibiting the flow separation over the surface. In addition, the increase of the jet momentum coefficient and the jet angle can both enhance the lift coefficient in the retreating side. Compared with a single jet, jet arrays have better control effects on improving the aerodynamic characteristic of rotors in forward flight.

Virtual engineering tools are not currently employed extensively during the certification and commissioning of flight simulator motion systems. Subjective opinion is regarded as sufficient for most applications, as it provides verification that the motion platform does not cause false cueing. However, the results of this practice are systems that may be far from optimal for their specific purpose. This paper presents a new method for tuning motion systems objectively using a novel tuning process and tools which can be applied throughout the simulators life-cycle. The use of the tuning method is shown for a number of simulated test cases.

Multiple Attribute Decision Analysis (MADA), known to be simple and convenient, is one of the most commonly used methods for Effectiveness Evaluation of Fighter (EEF), in which the attribute weight assignment plays a key role. Generally, there are two parts in the index system of MADA, i.e. performance index and decision index (or label), which denote the specific performance and the category of the object, respectively. In some index systems of EEF, the labels can be easily obtained, which are presented as the generations of fighters. However, the existing methods of attribute weight determination usually ignore or do not take full advantage of the supervisory function of labels. To make up for this deficiency, this paper develops an objective method based on fuzzy Bayes risk. In this method, a loss function model based on Gaussian kernel function is proposed to cope with the drawback that the loss function in Bayes risk is usually determined by experts. In order to evaluate the credibility of assigned weights, a longitudinal deviation and transverse residual correlation coefficient model is designed. Finally, a number of experiments, including the comparison experiments on University of California Irvine (UCI) data and EEF, are carried out to illustrate the superiority and applicability of the proposed method.

Optimisation of aircraft ground operations to reduce airport emissions can reduce resultant local air quality impacts. Single engine taxiing (SET), where only half of the installed number of engines are used for the majority of the taxi duration, offers the opportunity to reduce fuel consumption, and emissions of NOX, CO and HC. Using 3510 flight data records, this paper develops a model for SET operations and presents a case study of London Heathrow, where we show that SET is regularly implemented during taxi-in. The model predicts fuel consumption and pollutant emissions with greater accuracy than previous studies that used simplistic assumptions. Without SET during taxi-in, fuel consumption and pollutant emissions would increase by up to 50%. Reducing the time before SET is initiated to the 25th percentile of recorded values would reduce fuel consumption and pollutant emissions by 7–14%, respectively, relative to current operations. Future research should investigate the practicalities of reducing the time before SET initialisation so that additional benefits of reduced fuel loadings, which would decrease fuel consumption across the whole flight, can be achieved.

Speculation on near-term scientific reasons for the exploration of lunar pits is offered alongside comments on possible longer-term human exploitation. It is proposed that in order to determine whether or not one or more of the pits offer access the large subsurface voids e.g. a non-collapsed lava tube, a preliminary reconnaissance mission solely focused on obtaining lateral images (and/or LiDAR maps) is needed. Possible concept options for such a preliminary reconnaissance mission are discussed. It is suggested that one of the best possible strategies is to employ a micro-sized probe (~0.3 m) that would hop from a nearby main landing spacecraft to the selected pit. After the surface position of the main lander is determined accurately, the probe would perform a ballistic hop, or hover-traverse, a distance of ~3 km over the lunar surface using existing propulsive and guidance technology capability. Once hovering above the pit, the probe or a separate tethered imaging unit would then be lowered into the pit to acquire the necessary subsurface void topology data. This data would then be transmitted back to Earth, directly, via the lander, or via a store-and-forward orbiting relay. Preliminary estimates indicate that a probe of ~14 kg (dry mass) is viable using a conventional hydrazine monopropellant system with a propellant mass fraction of less than ~0.2 (20%) including margins, suggesting a piggyback architecture would be feasible.

Previously, the concept of Ply Drop Sequence (PDS) is introduced by the authors for the designing of composite laminated structures with multiple regions. Compared to deleting a contiguous innermost/outermost plies in the classical guide-based blending, using PDS is more flexible than dropping plies between adjacent regions. In this article, a new blending model called the Permutation for Panel Sequence (PPS) blending model is proposed to correct the problem of repeated searching of discrete points in the design space for the previous PDS blending model. The proposed method is also applied to an 18-panel horseshoe benchmark problem. The results demonstrate that the useful searching points in the PPS method are less than those in the PDS method when the number of the panels is less than the number of plies in the guide laminate, and the PPS method obtains a faster convergence speed compared with the PDS method.

In this paper, a non-linear tracking control algorithm is extended. The control objective of this research is to track a desired time-varying attitude of a satellite in the presence of inertia uncertainties and external disturbances, in order to be more suitable for implementation in a real-world application. In this investigation, the actuators are reaction wheels and the actuator dynamics are modelled in addition to the spacecraft dynamics. Thus, the control signal is DC motor voltage which is the most fundamental control variable and can be generated easily by a motor driver in practical cases. To achieve robust tracking of the desired time-varying attitude, a sliding mode controller is designed, and adaptive techniques are developed based on sliding mode control to overcome the inertia uncertainties and to estimate and compensate external disturbances. The kinematic equations of the satellite are expressed using quaternion parameters, and a novel control law will be derived by using a new facilitating approach in controller design, which is based on quaternion algebra, because of quaternion advantages, such as singularity rejection. Using this approach it will be more comfortable to deal with tedious mathematical operations, and on contrary with most of the previous studies, the terms corresponding to derivatives of the desired attitude are not neglected, and tracking capability is retained. The global stability of both methods (Sliding Mode Control (SMC) and adaptive sliding) is investigated using Lyapunov’s stability theorem. In order to validate the control methods, first, Simulink-ADAMS co-simulation of a 3-DOF attitude control is used to verify the algorithm performance and integrity, and finally, the control strategy is implemented on the Amirkabir University of Technology (AUT) 3-DOF attitude simulator for different types of non-linear attitudes. Both co-simulation and implementation results clearly illustrate the designed attitude control algorithm’s excellent performance in the various manoeuvres.

The primary objective of this research was to analyse the discrepancies in polymeric properties in Y12 aircraft tyres after 50 landings. Tensile strength and elongation test, abrasion test and hardness tests were carried out for the nose and main gear tyres after the completion of 50 landings. Surface morphologies of the tyre samples were observed using SEM. ISO specifications were followed for each experimental method during testing. There was a reduction in both median tensile strength and elongation at break in aircraft tyres after 50 landings but the reduction rate of both parameters were lower in Tyre 2 (nose wheel tyre) compared with Tyre 1 (left main wheel tyre) and Tyre 3 (right main wheel tyre). The highest percentage of mean volume loss was reported for Tyre 3 (3.88%). In addition, the least percentage of mean volume loss was obtained in Tyre 2 (2.98%). The percentage of hardness reduction was highest in Tyre 2 (6.9%). The surface roughness was induced to the tyre surface after completion of 50 landings.

We report on the flight test results of an Airbus 320 during novel advanced required navigation performance (RNP) procedures which contain a fixed radius turn that delivers the aircraft onto a short instrument landing system (ILS) precision final. Moreover, the advanced RNP part contains altitude constraints and/or a coded vertical path angle. The approaches were flown automatically with guidance and autothrust as computed by the flight management system. We investigated the use of the fixed radius in conjunction with vertical path options onto (a) the performance of the speed profile for arrival time optimisation, (b) the vertical path during the RNP part of the procedure and (c) the performance of the autoland capability after a curved transition onto an ILS.

For the trials, we designed five instrument approaches to a runway equipped with ILS. A RF curve terminates at the ILS intercept point at heights of 550, 750, 1000, 1500 and 2000 ft above aerodrome level and each approach had four different initial approach fixes which corresponded to a track angle change of 30°, 60°, 90° and 180° during the constant radius turn-to-final. We constructed the procedure such that the altitude constraints at initial, intermediate and final approach fix describe a continuous vertical path with minus 2° inclination, transitioning to the –3° glide path of the ILS and intercepting the glide path from below. In all cases, the land mode of the flight guidance computer became active between 316 and 381ft radar altitude. The vertical path following error depended on the coding of the procedure in the database. With coded vertical path angle and altitude constraints, the vertical path following error was never greater than +57 m (above desired flight path) during the RNP part when flown by the automatic flight guidance system without any pilot intervention.