A growing interest in constellations of small satellites has recently emerged due to the increasing capability of these platforms and their reduced time and cost of development. However, in the absence of dedicated launch services for these systems, alternative methods for the deployment of these constellations must be considered which can take advantage of the availability of secondary-payload launch opportunities. Furthermore, a means of exploring the effects and tradeoffs in corresponding system architectures is required. This paper presents a methodology to integrate the deployment of constellations of small satellites into the wider design process for these systems. Using a method of design-space exploration, enhanced understanding of the tradespace is supported , whilst identification of system designs for development is enabled by the application of an optimisation process. To demonstrate the method, a simplified analysis framework and a multiobjective genetic algorithm are implemented for three mission case-studies with differing application. The first two cases, modelled on existing constellations, indicate the benefits of design-space exploration, and possible savings which could be made in cost, system mass, or deployment time. The third case, based on a proposed Earth observation nanosatellite constellation, focuses on deployment following launch using a secondary-payload opportunity and demonstrates the breadth of feasible solutions which may not be considered if only point-designs are generated by a priori analysis. These results indicate that the presented method can support the development of future constellations of small satellites by improving the knowledge of different deployment strategies available during the early design phases and through enhanced exploration and identification of promising design alternatives.

RANS models remain an attractive turbulence simulation method which could provide some open jet aerofoil interaction analysis at a fraction of the cost of a high-fidelity LES approach. The present work explores the potential and limitations of RANS in this context by simulating an open jet aerofoil noise experiment using the aerospace oriented Menter SST RANS model. This model’s tendency to transition at a critical Reynolds number lower than the experimental value was found to impact the boundary layer development. However, the introduction of a low-Re correction improved the prediction of surface pressure and skin friction, enabling the suction surface separation bubble to be captured. The free shear layer’s virtual origin characteristics exhibited sensitivity to the interaction with the aerofoil, which can be developed into a metric of the interaction. The main challenge for RANS was accounting for the rise in background disturbance level in the working section, which is caused by the high-turbulence intensity in the free shear layers.

Advanced cooling techniques involving internal enhanced heat transfer and external film cooling and thermal barrier coatings (TBCs) are employed for gas turbine hot components to reduce metal temperatures and to extend their lifetime. A deeper understanding of the interaction mechanism of these thermal protection methods and the conjugate thermal behaviours of the turbine parts provides valuable guideline for the design stage. In this study, a conjugate heat transfer model of a turbine vane endwall with internal impingement and external film cooling is constructed to document the effects of TBCs on the overall cooling effectiveness using numerical simulations. Experiments on the same model with no TBCs are performed to validate the computational methods. Round and crater holes due to the inclusion of TBCs are investigated as well to address how film-cooling configurations affect the aero-thermal performance of the endwall. Results show that the TBCs have a profound effect in reducing the endwall metal temperatures for both cases. The TBC thermal protection for the endwall is shown to be more significant than the effect of increasing coolant mass flow rate. Although the crater holes have better film cooling performance than the traditional round holes, a slight decrement of overall cooling effectiveness is found for the crater configuration due to more endwall metal surfaces directly exposed to external mainstream flows. Energy loss coefficients at the vane passage exit show a relevant negative impact of adding TBCs on the cascade aerodynamic performance, particularly for the round hole case.

The maximisation of control power is considered for an aircraft with multiple control surfaces, with the force and moment coefficients specified by polynomials of the control surface deflections of degree two. The optimal deflections, which maximise and minimise any force or moment coefficient, are determined subject to constraints on the range of deflection of each control surface. The results are applied to a flying wing configuration to determine: (i/ii) the pitch trim, at the lowest drag for the fastest climb, and at the highest drag for the steepest descent; (iii) the maximum and minimum pitching moment; (iv) the maximum and minimum yaw control power and the fraction needed to compensate an outboard engine failure for several propulsion configurations; (v) the maximum and minimum rolling moment. The optimal use of all control surfaces has significant advantages over using just one, e.g. the range of drag modulation with pitch trim is much wider and the maximum and minimum available control moments larger.

Launching orbital and suborbital rockets from a high altitude is beneficial because of e.g. nozzle optimisation and reduced drag. Aircraft and gas balloons have been used for the purpose. Here we present a concept where a balloon is filled with pure water vapour on ground so that it rises to the launch altitude. The system resembles a gas balloon because no onboard energy source is carried, and no hard objects fall down. We simulate the ascent behaviour of the balloon. In the baseline simulation, we consider a 10 tonne rocket lifted to an altitude of 18 km.We model the trajectory of the balloon by taking into account steam adiabatic cooling, surface cooling, water condensation and balloon aerodynamic drag. The required steam mass proves to be only 1.4 times the mass of the rocket stage, and the ascent time is around 10 minutes. For small payloads, surface cooling increases the relative amount of steam needed, unless insulation is applied to the balloon skin. The ground-filled steam balloon seems to be an attractive and sustainable method of lifting payloads such as rockets into high altitude.

Numerical simulations of ship/rotor-coupled flowfield have been performed to investigate the rotational direction effects on a shipborne single-rotor helicopter in different deck landing trajectories (i.e., lateral and longitudinal translation) based on Reynolds-averaged Navier-Stokes (RANS) solver. Both the momentum source model and moving overset mesh model are employed to simulate the effect of the rotor on the ship airwake for different levels of fidelity requirement. The aerodynamic loading characteristics in terms of time-averaged and root-mean-square (RMS) thrust and pitch and roll moments are compared for two helicopter rotors with opposite rotation directions in a starboard 30 degrees wind condition. The time-averaged results show that the mean thrust of a counterclockwise rotor is greater than that of a clockwise rotor, particularly in the lateral translation phase. This suggests that a helicopter with a counterclockwise rotor could provide more collective control margin under this condition. Furthermore, a more significant reduction in pitch moment is experienced by the counterclockwise rotor during the two landing trajectories, and thus the effect of the aircraft being pulled towards the hangar tends to be more severe on the helicopter with the counterclockwise rotor. RMS loading results indicate that the unsteady loading levels on the clockwise rotor are much higher than that of the counterclockwise rotor in all three axes for most of the lateral and longitudinal translation phases. As a result, the pilot is likely to experience a higher workload when operating a helicopter with a clockwise rotor in the case of a deck landing in this wind condition.

Stiffened wing and fuselage panels often have a postbuckling reserve of strength, enabling them to carry loads far in excess of their critical buckling loads. Therefore allowing for postbuckling in design can reduce their weight, hence reducing fuel consumption and environmental impact. The present paper extends the postbuckling analysis in the exact strip software VICONOPT to more accurately reflect the skewed mode shapes arising from shear load and anisotropy. Such mode shapes are represented by a series of sinusoidal responses with different half-wavelengths which are coupled together using Lagrangian multipliers to enforce the boundary conditions. In postbuckling analysis the in-plane deflections involve responses with additional half-wavelengths which are absent from the out-of-plane deflection series. Numerical results are presented and compared with finite element analysis for validation. The present analysis gives close results compared to the finite element and finite strip methods and saves computational time significantly.

Route planning and airspace sectorisation are two central tasks in air traffic management.Traditionally, the routing and sectorisation problems were considered separately, with aircraft trajectories serving as input to the sectorisation problem and, reciprocally, sectors being part of the input to the path finding algorithms.

In this paper we propose a simultaneous design of routes and sectors for a transition airspace. We compare two approaches for this integrated design: one based on mixed integer programming, and one Voronoi-based model that separates potential “hotspots” of controller activity resulting from the terminal routes.

We apply our two approaches to the design of Stockholm Terminal Maneuvering Area.

An important prerequisite for the design, assessment and certification of aircraft and their associated control systems is a quantitative specification of the environment in which the aircraft is intended to operate, for example, atmospheric gust. Gust loads on aircraft may induce detrimental influences such as increased aerodynamic and structural loads, structural deformation and decreased flight dynamic performance. This paper presents a systematic and comprehensive overview of important concepts and applications of gust loads on aircraft. This overview includes a brief research background, concepts, research techniques, influences and load alleviation measures of gust. Finally, we summarise some potential improvements in the future work. It is also recommended to learn from previous experiences to avoid aviation accidents due to flight through atmospheric gusts and turbulence.

Hypersonic air-breathing propulsion can improve cost and flexibility of Low Earth Orbit (LEO) satellite launch missions. However, at the high flight Mach numbers required for access-to-space, performance margins are extremely tight. Techniques to improve mixing efficiency can push this technology forward. However, these are required to produce a minimal increase in losses and heat loads to be viable. The use of inlet-generated vortices in scramjets for mixing enhancement was previously studied. These vortices interact with the injected fuel plume, stretching it and increasing its effective surface for mixing. Moreover, these vortices are intrinsic to the flowfield. Therefore, contrary to other methods, when using inlet vortices mixing is enhanced without producing additional heat loads or losses. This work studies the vortex-injection interaction through numerical RANS simulations. A non-dimensional variable defining the quality of the plume shape for mixing purposes is proposed. This parameter is used to assess the effect of vortex intensity and injector location on fuel plume shape. The results show the ability of inlet vortices to modify fuel plume shape significantly increasing fuel mixing rate with minimal impact on losses.

With the aims of overcoming the limitations of the existing basic flow model derived from an axisymmetric generating body and extending the aerodynamic design method of the airframe/inlet integrated waverider vehicle, this study develops an upgraded basic flow model derived from an axisymmetric shock wave. It then upgrades the design method for airframe/inlet integration of an air-breathing hypersonic waverider vehicle, which is termed the ‘full-waverider vehicle’ in this study. In this paper, first, the design principle and method for the upgraded full-waverider vehicle derived from an axisymmetric basic shock wave are described in detail. Second, an upgraded basic flow model that accounts for both internal and external flows is derived from an axisymmetric basic shock wave by use of both the streamline tracing method and the method of characteristics (MOC). Third, the upgraded full-waverider vehicle is developed from the upgraded basic flow model by the streamline tracing method. Fourth, the design theories and methodologies of both the upgraded basic flow model and the upgraded full-waverider vehicle are validated by a numerical computation method. Finally, the aerodynamic performances and viscous effects of both the upgraded basic flow model and the upgraded full-waverider vehicle are analysed by numerical computation. The obtained results show that the upgraded basic flow model and aerodynamic design method are effective for the design of the airframe/inlet integration of an air-breathing hypersonic waverider vehicle.

Steps required for proper acquisition and processing of laser Doppler velocimetry data for turbomachinery research applications are addressed. Turbomachinery applications are difficult due to the small internal passages, high-frequency fluctuations, large turbulence intensities, and strong secondary flows resulting in low overall signal-to-noise ratios and narrowband noise sources that cannot be removed by simple band-pass filters. Special aspects that must be considered for successful and accurate laser Doppler velocimetry studies to be conducted in turbomachinery are discussed. Specifically, the design of the measurement volume size, reflection mitigation, engineering of seed particle size and injection schema, and alignment of the traverse mechanism are addressed in terms of their importance (from literature sources) and the solutions implemented by the authors. These techniques have been applied to successfully obtain three-component, unsteady velocity data in a high-speed centrifugal compressor for aeroengine application. Processing techniques are also presented including a novel mixture-model-based statistical method for narrowband noise isolation developed by the authors. The method, validation steps, and example results are presented, showing the successful rejection of noise with high accuracy, a low failure rate, and a significant reduction in required manual inspection. This newly developed method elucidated flow features that were not clear prior to the noise removal.

Experimental and numerical investigations into the linear and nonlinear aeroelastic behaviour of very flexible High Altitude Long Endurance (HALE) wings are conducted to assess the effect of geometrical nonlinearities on wings displaying moderate-to-large displacement. The study shows that the dynamic behaviour of wings under large deflection, and specifically the edgewise and torsion natural frequencies and modal characteristics, are largely affected by the presence of geometrical nonlinearities. A modular wing structure has been manufactured by rapid prototyping and it has been tested to characterise its dynamic and aeroelastic behaviour. At first, several simple isotropic cantilever beams with selected crosssections are numerically investigated to extract their modal characteristics. Experiments are subsequently conducted to validate the geometrically nonlinear dynamics behaviour due to high tip displacement and to understand the influence of the beam cross-section geometry. The structural dynamics and aeroelastic analysis of a very flexible modular selected wing is then investigated. Clean-wing wind-tunnel tests are carried out to assess flutter and dynamic response. The wind-tunnel model display interesting aeroelastic features including the substantial influence of the wing large deformation on its natural frequencies and modal characteristics.

The aim of this work is to develop a calculation model based on the method of characteristics making it possible to study the effect of the stagnation pressure of the combustion chamber on the 2D and axisymmetric minimum length nozzle design giving a uniform and parallel flow at the exit section. The model is based on the use of the real gas approach. The co-volume and the intermolecular interaction effect are taken into account by the use of the Berthelot state equation. The effect of molecular vibration is considered in our model to evaluate the behaviour of gas at a high temperature. In this case, the stagnation pressure and the stagnation temperature are important parameters in our model. The resolution of the algebraic equations is done by the finite difference corrector predictor algorithm. The validation of the results is controlled by the convergence of the critical section ratios calculated numerically as obtained by the theory. The mass and the thrust are evaluated to improve the efficiency of the nozzle. The comparison is made with the high temperature and perfect gas models. The application is made for air.

The model of Nano quad-rotors contains many uncertainties such as an external disturbance from a wind field, highly non-linear strong coupling between variables and body measurement errors. To deal with these uncertainties and control the Nano quad-rotors, a novel data-based disturbance observer (DO) is firstly proposed to observe disturbances from a wind field and perturbations from errors of parameter estimation. Then the DO is used to improve the conventional dynamic inversion (DI) method to obtain an enhanced dynamic inversion (EDI) method, which relies only on roughly estimated geometrical parameters, thus eliminating the largest flaw of conventional DI, namely depending on detailed plant information. Simulation results show that the method proposed achieved good trajectory tracking with only roughly estimated geometrical values under wind field; the DO proposed can accurately estimate disturbance from a wind field and perturbation from error of parameter estimation.

The takeoff-mass of a two-stage-to-orbit Rocket-Based Combined Cycle Engine-Rocket (RBCC-RKT) launch vehicle is a crucial factor in its comprehensive performance. This paper optimizes the takeoff-mass together with the trajectory by reformulating it to a nonlinear optimal control problem. The range of the second stage rocket mass is considered as a process constraint. When the scopes of initial and terminal states are specified, the problem can be solved by using the Gauss pseudo-spectral method (GPM). In order to reduce the convergent difficulty caused by using table data, the data in different stages are utilized by employing an integrated interpolation strategy through the optimization. Simulation results show that the mass can be effectively optimized to meet the inertia mass ratio constraint of the first-stage, and the separation of Mach number and altitude can be optimized at the same time.

Numerical simulations have been carried out for a 32.16-ft-diameter rotor in autorotational forward flight. Coupled flapping and rotational equations were solved using the transient simulation method (TSM) to ascertain the quasistatic torque equilibrium conditions. The Pitt/Peters inflow theory was adopted in the simulations, and an airfoil look-up table made by a compressible Navier-Stokes solver was used. The adverse cyclic and collective pitch inputs were introduced in a similar fashion to helicopter control in that the cyclic lever is pulled back and the collective lever is pushed down for increasing airspeeds. The simulation results showed that the longitudinal cyclic pitch input combined with a lowered collective pitch increases the rotating torque for a low shaft angle and an advance ratio greater than one, producing both high lift and a high lift-to-drag ratio. Upon introducing the adverse cyclic and collective pitch inputs, the control range broadened, and a torque equilibrium condition was detected at 414.7kt (700ft/s) of airspeed in the simulation.

In order to increase the speed, precision and robustness against the engine failure in solving optimal endo-atmospheric ascent trajectory of a launch vehicle, a rapid multi-layer solving method with improved numerical algorithms was proposed. The proposed method is capable of decomposing a large number of intervals into multiple layers with advantageous convergence property. Firstly, the problem of solving optimal endo-atmospheric ascent trajectory, which was subjected to path constraints and terminal constraints, was transformed into a Hamilton Two Point Boundary Value Problem (TPBVP). Then, through the finite difference method and numerical solving algorithm, the Hamilton TPBVP was iteratively solved with fewer initial discrete intervals. The initial values of higher-layer iterations were obtained by interpolating convergent solutions at sparse nodes into the doubly discrete nodes of high layers. The process was repeatedly performed until the solving precision met the requirements. To decrease the calculation load in solving TPBVPs, two improved solving algorithms without and with fewer Jacobian calculations were studied, respectively the Derivative-free Spectral Algorithm for Nonlinear Equations(DF-SANE) combined with the improved derivative-free nonmonotone line search strategy, and the Modified Newton method with a relaxation factor in combination with the Inverse Broyden Quasi-Newton method, denoted as ‘MN-IBQ’. Simulation verifications showed that the multi-layer method had significantly higher solving speed than the single-layer method. For the improved numerical algorithms, the DF-SANE was trapped in the local convergence problem. While using the proposed MN-IBQ can further increase the solving rate. Typical engine failure simulations showed that the multi-layer method with the MN-IBQ algorithm had not only significantly higher solving speed but also stronger robustness, where the traditional single-layer method could not adapt. In addition, the thrust loss tolerance limits for the multi-layer solving method were given for different engine failure times. The results show promising potential of the proposed approach in trajectory online generation and closed-loop guidance of launch vehicles at the endo-atmospheric ascent stage.