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.

System software subsystems in an unmanned aircraft system share hardware resources due to space, weight, and power constraints. Such subsystems have different criticality, requirements, and failure rates, and can cause undesired interference when sharing the same hardware. A component with high failure rate can reduce the reliability of the system unless a fault containment mechanism is adopted.

This work proposes an asymmetric multiprocessor architecture to establish isolation at the hardware level for distributed implementation of safety-critical subsystems along with user defined payload subsystems on the same hardware with minimally reduced reliability of the system. To achieve that, subsystems are strategically segregated in separate processors, connected to an on-chip protective interconnect for inter-processor communications. A custom watchdog and reset mechanism are implemented to reset a specific processor without affecting the entire system if required. The architecture is demonstrated on a FPGA chip. In addition, an example of an optimised distribution is provided for a specific flight control system with five subsystems.

The maximum attainable performance of small gas turbines represents a strong limitation to the operating altitude and endurance of high-altitude unmanned aerial vehicles (UAVs). Significant improvement of the cycle thermal efficiency can be achieved through the introduction of heat exchangers, with the consequent increase of the overall engine weight. Since semi-closed cycle engines can achieve a superior degree of compactness compared to their open cycle counterparts, their use can offset the additional weight of the heat exchangers. This paper applies semi-closed cycles to a high-altitude UAV propulsion system, with the objective of assessing the benefits introduced on the engine performance and weight. A detailed model has been created to account for component performance and size variation as function of thermodynamic parameters. The sizing has been coupled with a multi-objective optimisation algorithm for minimum specific fuel consumption and weight. Results of two different semi-closed cycle configurations are compared with equivalent state-of-the-art open cycles, represented by a recuperated and an intercooled-recuperated engine. The results show that, for a fixed design power output, engine weight is approximately halved compared to state-of-the-art open cycles, with slightly improved specific fuel consumption performance. Optimum semi-closed cycles furthermore operate at higher overall pressure ratios than open cycles and make use of recuperators with higher effectiveness as the mass penalty of the recuperator is smaller due to the lower engine mass flow rates.

This study is aimed at establishing a linear holding (LH) procedure instead of a conventional air holding stack to minimize the effects of airborne delays in terms of air traffic management and fuel consumption. This paper uses both actual flight data and the Base of Aircraft Database (BADA) model to obtain fuel consumption for level flight and descent segments, separately.

The total fuel savings obtained by using actual flight data (16%) and the BADA model (10%) indicate that the LH is found to be more advantageous compared to a conventional holding procedure. Furthermore, the recommended LH procedure could be a promising solution for keeping aircraft in a narrow area that could be considered to be an effective method for airspace usage.

The special application environment puts forward the higher requirement of reliability of parts made from titanium alloy Ti–6Al–4V, which is closely related to the machining-induced residual stress. For the fact of the non-linear distribution of residual stress beneath the machined surface, distribution of peripheral milling-induced residual stress and its effect on fatigue performance of titanium alloy Ti–6Al–4V are still confusing. In the present study, residual stress profile induced by peripheral milling of Ti–6Al–4V is first studied. And then, energy criteria are proposed to characterise the whole state of the residual stress field. Finally, the effects of residual stress profile and surface energy on tensile–tensile fatigue performance of titanium alloy Ti–6Al–4V are discussed. The conclusions were drawn that the variation trend of surface residual stress (σr,Sur), maximum compressive residual stress (σC,ax), location (hr0) and response depth (hry) of residual stress profile with cutting parameters showed a similar pattern for both measure directions those parallel (σ1) and perpendicular (σ3) to the cutting direction. Cutting speed and feed rate have a main effect on surface residual stress, and the depth of cut has little effect on all the four key factors of residual stress profile. With the increase of cutting speed and feed rate, machining-induced surface energy tends to become larger. But increasing the depth of cut caused the strain energy stored in unit time to decrease. Furthermore, the effect of depth of cut on surface energy was weakened when the value of cutting depth becomes larger. Both the surface compressive residual stress and the maximum compressive residual stress are beneficial for prolonging the fatigue life, while large value of machining-induced surface energy leads to a decrease of fatigue life. Analysis of variance result shows that maximum residual compressive stress has a greater impact on fatigue life than other residual stress factors.

Ornithopters and entomopters should be insensitive to the gusty environment during outdoor operations. Hence, it becomes imperative to understand their behaviour under the influence of gust for ensuring stable flight. In light of this, the present numerical study focused on understanding the aerodynamics of flapping wings with five different planform shapes under the influence of a spatiotemporally varying frontal gust. 3D, unsteady, laminar, and incompressible Navier-Stokes equations were solved using finite volume formulation. A canonical case of asymmetric 1 degree of freedom (DoF) flapping kinematics was considered. Horizontal and vertical force patterns in constant and gusty inflow conditions were numerically computed and compared. Findings were analyzed quantitatively by comparing the differences in the instantaneous force patterns, ordinal scoring approach, and phase space plots. Qualitative comparisons were made based on plots of vortex structures and surface pressure contours for constant and gusty inflow conditions for wings with different planform shapes. Spanwise Lagrangian Coherent Structures (LCS) of all the five wings were also compared. Studies revealed that the elliptical wing exhibited low sensitivity and inverse semi-elliptical wing exhibited high sensitivity to the gusty inflow. Rectangular, triangular and semi-elliptical shaped wings were moderately sensitive to the gusty inflow. This finding, within the limitations of the flapping kinematics and simulation conditions considered for the present study, supported the fact that many natural flyers like forest raptors, non-migratory passerines, pheasants, and partridges have adopted elliptical wing planform for efficient flight.

Previous studies demonstrated that laminar separation bubbles (LSBs) in the absence of external disturbances or forcing are intrinsically unstable with respect to a three-dimensional instability of centrifugal nature. This instability produces topological modifications of the recirculation region with the introduction of streamwise vorticity in an otherwise purely two-dimensional time-averaged flows. Concurrently, the existence of spanwise inhomogeneities in LSBs have been reported in experiments in which the amplification of convective instability waves dominates the physics. The co-existence of the two instability mechanisms is investigated herein by means of three-dimensional parabolised stability equations. The spanwise waviness of the LSB on account of the primary instability is found to modify the amplification of incoming disturbance waves in the linear regime, resulting in a remarkable enhancement of the amplitude growth and a three-dimensional arrangement of the disturbance waves in the aft portion of the bubble. Present findings suggest that the oblique transition scenario should be expected in LSBs dominated by the convective instability, unless high-amplitude disturbances are imposed.

This paper aims at contributing to a better understanding of the effect of Tyler–Sofrin Modes (TSMs) on forced vibration responses by analysing a 4.5-stage research axial compressor rig. The first part starts with a brief review of the involved physical mechanisms and necessary prerequisites for the generation of TSMs in multistage engines. This review is supported by unsteady CFD simulations of a quasi 2D section of the studied engine. It is shown that the amplitude increasing effect due to mistuning can be further amplified by the presence of TSMs. Furthermore, the sensitivity with respect to the structural coupling of the blades and the damping as well as the shape of the expected envelope is analysed.

The second part deals with the Rotor 2 blisk of the research compressor rig. The resonance of a higher blade mode with the engine order of the upstream stator is studied in two different flow conditions realised by different variable stator vane (VSV) schedules which allows to separate the influence of TSMs from the impact of mistuning. A subset of nominal system modes representation of the rotor is used to describe its mistuned vibration behaviour, and unsteady CFD simulations are used to characterise the present strength of the TSMs in the particular operating conditions. Measured maximum amplitude vs blade pattern and frequency response functions are compared against the predictions of the aeromechanical models in order to assess the strength of the TSMs as well as its influence on vibration levels.

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.

This paper presents the work carried out to evaluate the benefits and performance impacts of introducing a hydrogen fuel cell powered electric taxiing system to a conventional short-haul aircraft. Tasks carried out in this research and reported in this paper include the initial system design, hydrogen tank initial sizing, calculation of the impact on fuel burn and emissions and the evaluation of the effects on Direct Operating Cost (DOC). The Airbus A320 has been selected as the datum aircraft for sizing the system, and the benefits analysis is particularly focused on the fleet composition and financial data of a Europe-based, low-cost, large-scale A320 family operator in 2016. The maximum power capacity of 400 kW has been sized based on the rolling friction coefficient of 0.02. Based on the operator’s 2016 financial, up to 1% fuel reduction can be achieved using the proposed system and the reduction in total maintenance cost is expected to be up to 7.3%. Additionally, up to 5.97% net profit improvement is estimated in comparison with the annual after-tax profit of the datum operator in 2016.

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.

Experiments were carried out with air as the test gas to obtain the surface convective heating rate and surface pressure distribution on blunt and sharp cone models flying at hypersonic speeds. Tests were performed in a hypersonic shock tunnel at two different angles of attack: ${0}^\circ$ and ${5}^\circ$ with angles of rotation $\phi = {0}^\circ, {90}^\circ$, and ${180}^\circ$. The experiments were conducted at a stagnation enthalpy of 1.4MJ/kg, flow Mach number of 6.56 and free stream Reynolds number based on the model length of $9.1 \times {10}^{5}$. The effective test time of the shock tunnel is 3ms. The results obtained for cone model with a bluntness ratio of 0.2 were compared with sharp cone models for $\alpha ={0}^\circ$. The measured stagnation heat transfer value matched well with the theoretical value predicted by the Fay and Riddell correlation and with the CFD results.

This paper describes the extensive numerical investigation carried out on a 203-mm spin-stabilised projectile to study the effects of Magnus force at high angles of attack on the stability and flight-trajectory parameters, for further validation and incorporation in a 6-DOF trajectory solver for flight-stability analysis. Magnus force typically influences the course of flight by causing the projectile to drift from its intended path in addition to generation of inbuilt dynamic instabilities in pitch and yaw orientation and is a function of AoA and spin rate. This study is a consolidation of the authors’ previous research on the same caliber projectile but with time-accurate analysis. It has been found that typically, the Magnus force and moment calculation requires time-accurate Navier Stokes equations to be solved numerically for accurate prediction(1,2). Hence, to complete the extraction of static and dynamic coefficients derivatives, unstructured time-accurate CFD analysis on multiple configurations, ranging from subsonic to supersonic Mach regimes, has been evaluated using Large Eddy Simulation (LES) and found to be suitable for capturing the desired effects. However, the LES simulation requires non-dimensional wall distance (y+) of the order of 0.5 – 1, with LES_IQ > 75% thus, is computationally-intensive. In addition, to cover the entire flight envelope from Charge 1 (249 m/s) to Charge 7 (595 m/s), at spin rate from 500 rad/s to 750 rad/s, 30 cases have been evaluated to generate the time-accurate coefficient library for integration with 6-DOF solver analysis. The results obtained have been compared with the available experimental data and found to be in reasonable agreement. The results of 6-DOF solver, incorporating the extracted coefficients, were compared with firing-table results, which further validated the computational methodology. This study provides an insight on how opposite flow interacts with the attached boundary layer due to spin rate and generates a turbulent interacting flow with variation in vortical structures for Q-Criterion vortex-flow visualization.

In this paper, a Gaussian process regression (GPR)-based novel method is proposed for non-linear aerodynamic modelling of the aircraft using flight data. This data-driven regression approach uses the kernel-based probabilistic model to predict the non-linearity. The efficacy of this method is examined and validated by estimating force and moment coefficients using research aircraft flight data. Estimated coefficients of aerodynamic force and moment using GPR method are compared with the estimated coefficients using maximum-likelihood estimation (MLE) method. Estimated coefficients from the GPR method are statistically analysed and found to be at par with estimated coefficients from MLE, which is popularly used as a conventional method. GPR approach does not require to solve the complex equations of motion. GPR further can be directed for the generalised applications in the area of aeroelasticity, load estimation, and optimisation.

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.

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.

Generally, earth rotating and non-spherical perturbation of the earth in re-entry motion model are simplified using the standard trajectory guidance method. The re-entry motion is also simplified to horizontal motion and vertical motion and controlled, respectively. The simplification of re-entry motion model will lead to loss of motion accuracy and location accuracy. The direct decomposition will lead to the reduction of control accuracy because the horizontal motion and the vertical motion are coupled in re-entry. To improve the standard trajectory guidance method, the standard trajectory guidance method based on decoupling control variables and waypoint is proposed in this paper. The proposed guidance method will not simplify earth rotating and non-spherical perturbation of the earth in motion equation or decompose the re-entry motion to horizontal motion and vertical motion. Trajectory waypoint is adopted to reduce the change frequency of tracking states, because tracking states change frequently if the entire standard trajectory is tracked in real time.

In order to prepare an advanced 4-stage high-pressure compressor rig test campaign, details regarding both accomplishment and analysis of preliminary experiments are provided in this paper. The superior objective of the research project is to contribute to a reliable but simultaneously less conservative design of future high pressure blade integrated disks (blisk). It is planned to achieve trend-setting advances based on a close combination of both numerical and experimental analyses. The analyses are focused on the second rotor of this research compressor, which is the only one being manufactured as blisk. The comprehensive test program is addressing both surge and forced response analyses e.g. caused by low engine order excitation. Among others the interaction of aeroelastics and blade mistuning is demanding attention in this regard. That is why structural models are needed, allowing for an accurate forced response prediction close to reality. Furthermore, these models are required to support the assessment of blade tip timing (BTT) data gathered in the rig tests and strain gauge (s/g) data as well. To gain the maximum information regarding the correlation between BTT data, s/g-data and pressure gauge data, every blade of the second stage rotor (28 blades) is applied with s/g. However, it is well known that s/g on blades can contribute additional mistuning that had to be considered upon updating structural models.

Due to the relevance of mistuning, efforts are made for its accurate experimental determination. Blade-by-blade impact tests according to a patented approach are used for this purpose. From the research point of view, it is most interesting to determine both the effect s/g-instrumentation and assembling the compressor stages on blade frequency mistuning. That is why experimental mistuning tests carried out immediately after manufacturing the blisk are repeated twice, namely, after s/g instrumentation and after assembling. To complete the pre-test program, the pure mechanical damping and modal damping ratios dependent on the ambient pressure are experimentally determined inside a pressure vessel. Subsequently the mistuning data gained before is used for updating subset of nominal system mode (SNM) models. Aerodynamic influence coefficients (AICs) are implemented to take aeroelastic interaction into account for forced response analyses. Within a comparison of different models, it is shown for the fundamental flap mode (1F) that the s/g instrumentation significantly affects the forced response, whereas the impact of assembling the compressor plays a minor role.

Data mining is a process of finding correlations and collecting and analysing a huge amount of data in a database to discover patterns or relationships. Flight delay creates significant problems in the present aviation system. Data mining techniques are desired for analysing the performance in which micro-level causes propagate to make system-level patterns of delay. Analysing flight delays is very difficult – both when looking from a historical view as well as when estimating delays with forecast demand. This paper proposes using Decision Tree (DT), Support Vector Machine (SVM), Naive Bayesian (NB), K-nearest neighbour (KNN) and Artificial Neural Network (ANN) to study and analyse delays among aircrafts. The performance of different data mining methods is found in the different regions of the updated datasets on these classifiers. Finally, the result shows a significant variation in the performance of different data mining methods and feature selection for this problem. This paper aims to deal with how data mining techniques can be used to understand difficult aircraft system delays in aviation. Our aim is to develop a classification model for studying and reducing delay using different data mining methods and, in this manner, to show that DT has a greater classification accuracy. The different feature selectors are used in this study in order to reduce the number of initial attributes. Our results clearly demonstrate the value of DT for analysing and visualising how system-level effects happen from subsystem-level causes.

This paper addresses a methodology to parametrically size thermal control subsystems for high-speed transportation systems during the conceptual design phase. This methodology should be sufficiently general to be exploited for the derivation of Estimation Relationships (ERs) for geometrically sizing characteristics as well as mass, volume and power budgets both for active (turbopumps, turbines and compressors) and passive components (heat exchangers, tanks and pipes). Following this approach, ad-hoc semi-empirical models relating the geometrical sizing, mass, volume and power features of each component to the operating conditions have been derived. As a specific case, a semi-empirical parametric model for turbopumps sizing is derived. In addition, the Thermal and Energy Management Subsystem (TEMS) for the LAPCAT MR2 vehicle is used as an example of a highly integrated multifunctional subsystem. The TEMS is based on the exploitation of liquid hydrogen boil-off in the cryogenic tanks generated by the heat load penetrating the aeroshell throughout the point-to-point hypersonic mission. Eventually, specific comments about the results will be provided together with suggestions for future improvements.