At the preliminary design phase of a gas turbine, time is crucial in capturing new business opportunities. In order to minimise the design time, the concept of Preliminary Multi-Disciplinary Optimisation (PMDO) was used to create parametric models, geometry and cooling flow correlations towards a new design process for turbine housing and shroud segments. First, dedicated parametric models were created because of their reusability and versatility. Their ease of use compared to non-parameterised models allows more design iterations and reduces set-up and design time. A user interface was developed to interact with the parametric models and improve the design time. Second, geometry correlations were created to minimise the number of parameters used in turbine housing and shroud segment design. Third, a correlation study was conducted to minimise the number of engine parameters required in cooling flow predictions. The parametric models, the geometry correlations, and the user interface resulted in a time saving of 50% and an increase in accuracy of 56% compared to the existing design system. For the cooling flow correlations, the number of engine parameters was reduced by a factor of 6 to create a simplified prediction model and hence a faster shroud segment selection process.

Icing in conditions where clouds contain both liquid and solid phase particles has attracted considerable interest in recent years due to numerous in-flight incidents including engine rollbacks in the vicinity of deep convective clouds in tropical regions. These incidents have prompted certification authorities to investigate and extend the icing conditions to include solid and mixed-phase clouds for airworthiness certification. These efforts have resulted in the amendments issued by the Federal Aviation Administration (FAA) and European Aviation Safety Agency (EASA) to the certification specifications of large aircraft, FAR-25 and CS-25, respectively. Flight tests, laboratory tests and computer simulations are among the acceptable means to show compliance with these specifications. Considerable effort has been spent worldwide in order to develop icing simulation software for liquid phase clouds in the past four decades, but until recently, most of these software did not have the capability for solid- or mixed-phase clouds. One of the aims of the High Altitude Ice Crystals (HAIC) project funded by the European Commission within Framework Program 7 is to address this shortcoming. The present study combines the models related to solid- and mixed-phase icing that is developed within HAIC with an in-house numerical tool. The tool has four modules; a flow-field solution module that uses the Hess-Smith panel method, a module for computing droplet and ice crystal trajectories and collection efficiencies using the Lagrangian approach, a thermodynamic module, and an ice accretion module that utilises the Extended Messinger Model. The numerical tool is tested against experimental test cases including liquid and mixed-phase conditions for various aerofoil and axisymmetric intake geometries. The agreement of the obtained results with experimental data is encouraging.

Information regarding deformations in large and complex systems is necessary in the prediction of structural failures caused by un-natural flexural occurrences. Sensing systems which are used to predict shapes, in order to develop a global surface picture require high precision and lower time lag. In this work, a unique bio-inspired training mechanism for support vector regression is presented for shape sensing in structures mounted with Fiber Bragg Gratings. Experimental validation was carried out on a simply supported beam, loaded at different positions and an aircraft wing model for different types of bending. The resulting deflections at specified locations along the length of the beam and on both surfaces of the wing were interpreted from the wavelength shifts of the corresponding Fiber Bragg Gratings through the specially modified Support Vector Regression. The method has shown high accuracy, low computational requirements and enhanced prediction times. The proposed bio-inspired training method has also been compared with two conventional training methodologies.

A novel vortex generator design positioned upstream of a normal shock followed by a subsequent diffuser was investigated using large eddy simulations. In particular, “ramped-vane” flow control devices with three different heights relative to the incoming boundary layer thickness (0.34δ, 0.52δ and 0·75δ) were placed in a supersonic boundary layer with a freestream Mach number of 1.3 and a Reynolds number of 2400 based on the momentum thickness. This is the first numerical study to investigate the size effect of the ramped-vane for flow control device in terms of shape factor, flow separation and flow unsteadiness. The results showed that these devices generated strong streamwise vortices that entrained high-momentum fluid to the near-wall region and increased turbulent mixing. The devices also decreased shock-induced flow separation, which resulted in a higher downstream skin friction in the diffuser. In general, the largest ramped-vane (0.75δ) produced the largest reductions in flow separation, shape factor and overall unsteadiness. These results and a careful review of the literature study also determined the quantitative correlation of optimum VG height with Mach number, whereby h/δ~1 is often optimum for incompressible flows while higher Mach numbers lead to small optimum heights, tending towards h/δ~0.45 at M=2.5.

The scramjet is a rather a new technology and there are many issues related to their operation, especially when it comes to the combustion processes. Combustion in high-speed flows causes various problems such as flame instability and poor fuel–air mixing efficiency. One of the methods used to overcome these problems is to recess a cavity in the combustor wall where a secondary flow is generated. In this study, a computational fluid dynamics (CFD) code is developed to analyse the reacting flow passing through the cavity-based scramjet combustor. The developed code is based on three-dimensional coupled Navier–Stokes and finite rate chemistry equations. An ethylene-air reduced chemical reaction model is used as a fuel–air combination. The Spalart–Allmaras model is utilised for turbulence closure. The non-dimensional form of the flow and chemical reaction equations are discretised using a finite volume method. The Jacobian-Free Newton–Krylov (JFNK) method is used to solve the coupled system of non-linear equations. The JFNK is a matrix-free solution method which improves the computational cost of Newton’s method. The parameters that affect the performance of the JFNK method are studied in the analysis of a scramjet combustor. The influence of the forcing term on the convergence of the JFNK method is studied in the analysis of scramjet combustor. Different upwind flux vector splitting methods are utilised. Various flux limiter techniques are employed for the calculations of higher order flux vectors. The effects of flux vector splitting and flux limiter methods on the convergence and accuracy of the JFNK method are evaluated. Moreover, the variations of the mixing efficiency with fuel injection angles are studied.

Vertical profile inefficiency caused by low-level flight segments during arrival and departure operations has a great impact on the financial cost of operators as well as negative environmental effects. In this regard, continuous descent and climb operations for arrival and departure phases are mostly preferred as one of the methods that reduce or eliminate vertical profile inefficiencies. In this study, focus is put on the arrival phase and the purpose is: (i) to analyse the vertical profiles of arrival traffic at Mugla Milas-Bodrum Airport (BJV), Antalya Airport (AYT) and Ankara Esenboga Airport (ESB) and to characterise the level-offs that occur in the arrival phase and (ii) to quantify the potential time, fuel and emission savings in the event of a shift of inefficient low-level flight segments to the cruise segment of equal distance. For the purpose of this paper, real-time flight data is used for the analysis of the vertical profiles of flights. Flight parameters, such as latitude, longitude, altitude, speed and so on are used to determine, visualise and characterise level-off segments. The Boeing Fuel Flow Method 2 is used to calculate emissions at any altitude. Analyses of the results show that 80%, 74% and 69% of arrival traffic at BJV, AYT and ESB, respectively. were exposed to at least one level-off lasting longer than 20 seconds and significant potential for time, fuel and emissions savings exist for three airports.

The present study investigates the behaviour of the shock train in a typical Ramjet engine under the influence of shock and expansion waves at the entry of a low aspect ratio (1:0.75) rectangular duct/isolator at supersonic Mach number (M = 1.7). The start/unstart characteristics are investigated through steady/unsteady pressure measurements under different back and dynamic pressures while the shock train dynamics are captured through instantaneous Schlieren flow visualisation. Two parameters, namely pressure recovery and the pressure gradient, is derived to assess the duct/isolator performance. For a given back pressure, with maximum blockage (9% above nominal), the duct/isolator flow is established when the dynamic pressure is increased by 23.5%. The unsteady pressure measurements indicate different scales of eddies above 80 Hz (with and without flap deflection). Under the no flap deflection (no back pressure) condition, the maximum fluctuating pressure component is 0.01% and 0.1% of the stagnation pressure at X/L = 0.03 (close to the entry of the duct) and X/L = 0.53 (middle of the duct), respectively. Once the flap is deflected (δ = 8°), decay in eddies by one order is noticed. Further increase in back pressure (δ ≥ 11°) leads the flow to unstart where eddies are observed to be disappeared.

Following the B777 accident at Heathrow in 2008, the certification authorities required Boeing, Airbus, and Rolls-Royce to conduct icing analysis and tests of their Rolls-Royce Trent engined aircraft fuel systems. The experience and the test data gained from these activities were distilled and released by Airbus to the EASA ICAR project for research and analysis. This paper provided an overview of the Airbus ice accretion and release tests. Brief narratives on the test rigs, the test procedure and methodology were given and key findings from the ice release investigations were presented. The accreted ice thickness was non-uniform; however, it is found typically c. $\mathrm{2\;\mathrm{m}\mathrm{m}}$ thick. Analysis of the accreted ice collected from the rig tests showed the ice was very porous. The porosity is very much dependant on how the water was introduced and mixed in the icing test rigs. The standard Airbus method produced accreted ice of higher porosity compared to that produced by the injection method. The porosity of the accreted ice from Airbus icing investigations was found to be c. 0.90. The relationship of permeability with porosity was inferred from published data and models for freshly fallen snow in the atmosphere. Derived permeability $\mathrm{7.0\times 10^{-9}\;\mathrm{\mathrm{\mathrm{m}}^{\mathrm{2}}}}$ was then applied in the CFD analysis of pipe flow with a porous wall lining to determine the shear stress on the accreted ice. It showed that 25%, 50% and 75% of the accreted ice has interface shear strength of less than $\mathrm{15.3\;\mathrm{Pa}}$, $\mathrm{20.7\;\mathrm{Pa}}$ and $\mathrm{26.1\;\mathrm{Pa}}$, respectively.

Virtual Engineering (VE), also known as Model-Based Systems Engineering (MBSE), is necessary in both current operational engineering qualifications and to help reduce the costs of future vertical lift design and analysis. As computational power continues to provide increasing capability to the rotorcraft engineering community to perform simulations in both real time and off line, it is imperative that the community develop verification and validation protocols and processes to certify these methods so that they can be reliably used to help reduce engineering cost and schedule. Computational Fluid Dynamics (CFD) has become a major Computational Science and Engineering (CSE) tool in the fixed wing and vertical lift communities, but it has not been developed to the point where it is accepted as a replacement for testing in certification of new or existing systems or vehicles. Since the rise of modern CFD in the 1980s, the promise of CFD’s capabilities has been met or exceeded, but its role in certification arguably remains less prominent than projected. The ability to implement transformative technologies further drives the need for CFD in design. To meet CFD’s role in certification, several goals must be met to provide a true “numerical experiment” from which accuracies (error estimates), sensitivities, and consistent application results can be extracted. This paper discusses the progress and direction towards developing CFD strategies for certification.

The next generation of civil large aero-engines will employ greater bypass ratios compared with contemporary architectures. This results in higher exchange rates between exhaust performance and specific fuel consumption (SFC). Concurrently, the aerodynamic design of the exhaust is expected to play a key role in the success of future turbofans. This paper presents the development of a computational framework for the aerodynamic design of separate-jet exhaust systems for civil aero-engines. A mathematical approach is synthesised based on class-shape transformation (CST) functions for the parametric geometry definition of gas-turbine exhaust components such as annular ducts and nozzles. This geometry formulation is coupled with an automated viscous and compressible flow solution method and a cost-effective design space exploration (DSE) approach. The framework is deployed to optimise the performance of a separate-jet exhaust for very-high-bypass ratio (VHBR) turbofan engine. The optimisations carried out suggest the potential to increase the engine’s net propulsive force compared with a baseline architecture, through optimum exhaust re-design. The proposed method is able to identify and alleviate adverse flow-features that may deteriorate the aerodynamic behaviour of the exhaust system.

This paper aims to build the finite element model of the composite sinusoidal specimens and to carry out the parametric analysis. In this paper, the damage behaviour and the energy-absorbing results of composite sinusoidal specimens have been studied by quasi-static crushing experiments. The failure mechanisms of specimens under quasi-static crushing is further analysed. A numerical simulation has been performed by using the finite element model code LS-DYNA. The numerical results, in terms of load -displacement data, have been compared against experimental data, and good agreement has been found. Moreover, a sensitivity study has been carried out by varying material properties in order to assess their influence on the numerical results, and the material parameter selection scheme is optimised based on the constructed corresponding response surfaces. The results show that the response surface model has passed the test of goodness of fit, and the optimisation method can effectively assist the finite element modelling, and greatly decrease the numbers of trial and error.

One of the key problems in the development of morphing aircraft is the morphing structure, which should be able to carry loads and change its geometry simultaneously. This paper investigates a compliant structure, which has the potential to change the dihedral angle of morphing wing-tip devices. The compliant structure is able to induce deformation by unsymmetrical stiffness allocation and carry aerodynamic loads if the total stiffness of the structure is sufficient.

The concept has been introduced by building a simplified model of the structure and deriving the analytical equations. However, a properly designed stiffness asymmetry, which is optimised, can help to achieve the same deformation with a reduced actuation force.

In this paper, round corrugated panels are used in the compliant structure and the stiffness asymmetry is introduced by changing the geometry of the corrugation panel. A new equivalent model of the round corrugated panel is developed, which takes the axial and bending coupling of the corrugated panel into account. The stiffness matrix of the corrugated panel is obtained using the equivalent model, and then the deflections of the compliant structure can be calculated. The results are compared to those from detailed finite element models built in the commercial software Abaqus. Samples with different geometries were manufactured for experimental tests.

After verifying the equivalent model, optimisation is performed to find the optimum geometries of the compliant structures. The actuation force of a single compliant structure is first optimised, and then the optimisation is performed for a compliant structure consisting of multiple units. A case study is used to show the performance improvement obtained.

In Japan, the working group for the Mars Exploration Aircraft continues to research and develop a Mars aircraft aiming to a future survey mission. To ensure the success of the flight mission, a self-localisation system with low-computational complexity is necessary. In the present research, a new self-localisation method is proposed using multispectral images. The algorithm is based on the simple mapping of image moment invariants and gradients calculated from several images at different wavelengths. The numerical simulations revealed the sufficient robustness of the proposed method to image noise. The estimation accuracy can be improved by increasing the number of the spectral images.

Unlike conventional ballistic parachutes, gliding parachutes have been extensively used as guided precision aerial delivery system (GPADS) platforms in recent years. The reasoning may be found in gliding and manoeuvering capabilities, which make this kind of ram-air parachutes superior for precision aerial delivery application. In contrast, wing-shaped configuration along with more design variables create a cumbersome design procedure for this type of parachute. Especially, when an affordable configuration is demanded, the design procedure will be a more important problem. In this respect, an innovative integrated design framework is proposed in which significant design aspects are considered so that the gliding parachute configuration can be optimiseoptimised through a bi-objective optimisation problem to find optimum cost for achievable gliding ranges. To do so, the configuration is defined with minimum required parameters and design space is constrained by performance and stability as significant design requirements to guarantee the feasibility of the solutions in practice. The objective functions are defined in terms of gliding characteristics and amount of materials for fabrication which are representative for maximum reachable stand-off distance and unit cost, respectively. As an effective numerical optimisation method, a niched multi-objective genetic algorithm (MOGA) is used to generate a pareto-optimal set for a specific payload mass whereas constraints are handled through a tournament selection process. Based on results, the provided pareto front can aid the designers in decision-making and trade-off between demanded objectives for a payload weight. The underlying design problem is solved using an all-at-once (AAO) approach and the design loop converges to favourable results in a reasonable time. Finally, as a comparative study, an optimiseoptimised affordable cargo parachute is proposed for the GPADS application in which the material cost is reduced by at least 25% with respect to an available low-cost gliding parachute with the same glide ratio.

The purpose of this paper is to outline the structure of the DLR integrated rotorcraft design process. The complexity of rotorcraft design requires the development of the tools directly by the specialists of the respective institutes, where the tools are continuously refined and published to authorised users. The integration of the tools into a suitable software framework by means of distributed computation and the harmonisation of the tools among each other are presented. This framework delivers a high level of modularity making the layout and testing of the process very flexible. This design environment covers the conceptual and preliminary design phases. Not only conventional main/tail rotor configurations can be designed, but also some other configurations with more than one main rotor. The fundamental concept behind the layout of the tools is demonstrated, especially the use of scaling and optimisation loops in connection with the different levels of fidelity and the different phases of design.

A nonlinear station-keeping control method for a multi-vectored propeller airship under unknown wind field with thrust saturation is developed, which is composed of three modules: nonlinear model predictive controller (NMPC), disturbance observer (DOB) and tracking differentiator (TD). The nonlinear kinematics and dynamics models are introduced, and the wind effect is considered by the wind-induced aerodynamic force. Based on both models, an explicit NMPC is designed. Then a nonlinear DOB is introduced to estimate the wind disturbance. A TD, showing the relationship between the maximum propulsion force and the maximum flight acceleration, is proposed to handle the thrusts’ amplitude saturation. Stability analysis shows that the closed-loop system is globally asymptotically stable. Simulations for a multi-vectored propeller airship are conducted to demonstrate the robustness and effectiveness of the proposed method.

The paper presents a laser guided bomb guidance law based on the linear quadratic differential game theory, where a case of two perpendicular planes with two state variables in each plane has been considered. The Kalman filtering method has been used for noise removal from the measured signals and for estimation of the missing state variable values needed for the optimal guidance law. Optimisation has been conducted with respect to minimisation of the performance index. Comparative analysis of different guidance laws is done. A statistical analysis is performed to obtain the terminal miss distance in dependence on total flight time.

Detailed numerical simulations have been carried out on a spiked blunt body with multiple hemispherical disks using a commercial CFD code in order to investigate their effectiveness in reducing the aerodynamic drag and heating. The base configuration is a hemispherical cylinder whose diameter is 40 mm with an overall length of 70 mm. The lengths of the aerospikes investigated are 1, 1.5, 2 and 2.5 times the base diameter of the cylinder and the radii of the aerodisks are varied between 0.05, 0.1, 0.15 and 0.2 times the diameter of the cylinder. Besides these, the position of the aerodisks is varied with the rearmost aerodisk placed at 25%, 50% and 75% along the length of the aerospike and the intermediate aerodisk for three-disk cases, positioned at 25%, 50% and 75% of the distance between the front and the rearmost disk. All the investigations have carried out at a freestream Mach number of 6.2 and Reynolds number of 2.64 × 107/m. It has been observed that the multidisk spikes are advantageous for the purpose of reduction of both aerodynamic drag and heating at hypersonic speed. The two aerodisk spiked configurations show better results in terms of aerodynamic heating and drag in comparison to the single-disk aerospikes while the three-disk spikes yield only a marginal reduction in aerodynamic drag over the two-disk configurations. For reduction of heat fluxes and heat transfer rates though, the three-disk configurations are extremely advantageous and give much larger reductions are compared to the two-disk configurations.

Charge Exchange (CEX) ion is the main factor causing the plume pollution. The distribution of CEX ions is determined by the distribution of beam ions and neutral atoms. Hence, the primary problem in the study of the plume is how to accurately simulate the distribution of beam ions and neutral atoms. At present, the most commonly used model utilised for the plume simulation is the analytical model proposed by Roy for the plume simulation of the NASA Solar Technology Application Readiness (NSTAR) ion thruster. However, this analytical model can only be applied to the ion beam with small divergence angles. In addition, the analytical model is no longer applicable to the simulation for the plume of a new type of ion thruster that appeared recently, which is called the annular ion thruster. In this paper, a 3D particle model is proposed for the plume simulation of ion thrusters consisting of the particle model for beam ions, the Direct Simulation Monte Carlo (DSMC) model for neutral atoms and the Immersed Finite Element-Particle In Cell-Monte Carlo Collision (IFE-PIC-MCC) model for CEX ions. Then, the plume of the NSTAR ion thruster is simulated by both Roy's model and the 3D particle model. The simulation results of both models are then compared with the experimental results. It is shown that the numerical results of the 3D particle model agree well with those of the analytical model and the experimental data. And this 3D particle model can also be used for other electric thrusters.

As propeller-driven aircraft are the best choice for short/middle-haul flights but their acoustic emissions may require improvements to comply with future noise certification standards, this work aims to numerically evaluate the acoustics of different modern propeller designs. Overall sound pressure level and noise spectra of various blade geometries and hub configurations are compared on a surface representing the exterior fuselage of a typical large turboprop aircraft. Interior cabin noise is also evaluated using the transfer function of a Fokker 50 aircraft. A blade design operating at lower RPM and with the span-wise loading moved inboard is shown to be significantly quieter without severe performance penalties. The employed Computational Fluid Dynamics (CFD) method is able to reproduce the tonal content of all blades and its dependence on hub and blade design features.