In this paper an experiment is described in which a detailed study was made of the flow field associated with a slender sharp-edged delta wing which was rolling steadily at zero angle of attack to an air stream. The investigation was made by performing two pressure surveys: first , one of the static pressure acting on the wing’s surface and second by measuring the total-head distribution in the neighbourhood of the wing. From the former the local rolling-moment coefficients, Clp, are evaluated and these are compared with the predictions for attached flow, thus assessing the contributions to the forces acting on the wing which arise as a consequence of the leading-edge separations. The second set of surveys is used to construct a picture of the flow-field details and this is compared with that known to occur on a similar wing when it is set at an angle of attack to the airstream. One interesting finding is that the secondary separation which appears to cause the discrepancy between the theoretical predictions and the measurements made on slender wings at incidence, is absent in this configuration and thus it is concluded that these data could be used for a more meaningful test of the theory.

By employment of the analytical method of characteristics and of a limiting procedure suitable for dealing with the trailing edge expansion, the influence of near-field flow on the far-field wave formation has been investigated for an incident flat delta wing with supersonic leading edges. Though confined in its scope to the front shock in the vertical plane of symmetry of the wing and to a homogeneous atmosphere without density and temperature gradients, the present analysis reveals features of flow which are interesting from the standpoint of the general theory of three-dimensional supersonic flow. It is found that the front shock due to a delta wing will as a rule be cancelled at a finite distance from the wing by the plane-wave expansion emanating from the trailing edge. The over-expansion must then give rise to a rear shock separate from the front one. Thus, at least in the plane of symmetry, a sharp-front wave signature can not, in general, be expected from the wing at a distance beyond the terminating point of the front shock. The boom signature then will be qualitatively different from that of a body of revolution. The general non-equivalence of a wing to a body of revolution in this respect should evoke some rethinking about sonic boom prediction and alleviation.

The three-dimensional motion of a body towed by a flexible cable is studied. An approximate solution is developed for the non-linear equations of motion of the system to determine cable shapes. This solution is then used to compute natural frequencies and damping rates for longitudinal and lateral motions of the system. Conditions necessary for the uncoupling of the lateral and longitudinal motions are defined.

When fluid flows in a tube which rotates about an arbitrary axis, the presence of centripetal and Coriolis acceleration components modify the velocity and pressure fields which exist in the absence of rotation. Barua considered the case of an incompressible fluid flowing in laminar motion through a cylindrical tube which was rotating about an axis perpendicular to itself with uniform angular velocity. For distances well away from the tube entrance Barua illustrated that secondary flow in the r-θ plane occurred and that the axial velocity profile was distorted towards the leading edge of the tube. Since the pressure gradient along the tube is proportional to the gradient of the axial velocity profile at the tube wall the rotation thus has a consequential influence on the resistance to flow offered by the tube.

Composite materials have been increasingly used in aircraft structures. However, these composite structures are susceptible to damage from external low-velocity impacts. In this paper, an impact identification algorithm is proposed to estimate the impact location and force time history simultaneously. A localisation method based on basis vectors is proposed, and the impact force time history is reconstructed by simplified transfer functions. The basis vector stands for the relationship between the impact location and the sensor signals, and the transfer function shows the relationship of the sensor signal and the force time history. An experiment is conducted on a flat glass fibre-epoxy matrix composite plate to verify the developed algorithm using only four sensors. The soft impactor and hard impactor are two typical impactors for impact events; therefore, the impact experiment is performed by the rubber and the steel impactors, respectively. The experimental results indicate that the proposed algorithm is feasible for the identification of impact events on plate-like composite structures.

The problem of the buckling under bi-axial compression is considered for flat rectangular isotropic plates with simple edge conditions and no lateral restraint. The buckling stress coefficient is plotted against the side ratio for various conditions of edge restraint and a known compression or tension on the sides of the plate. It is found that there is some reduction in the buckling stress if the sides of the plate are subjected to compressive stresses and, conversely, there is an increase for tensile stresses. Furthermore, for plates with large side ratios, there is a rapid decrease of the spanwise buckling stress as the chordwise compressive stress approaches its appropriate Euler instability value.

The recently developed theory of modal control has so far been mainly applied to the design of process control systems. This theory is, however, capable of much wider application and it is the principal object of this paper to illustrate its use as a basis for the synthesis of aircraft control systems. The particular illustrative problem considered is that of synthesising a lateral autopilot for bank-angle and roll-rate mode control. Typical aerodynamic data for an aircraft flying at M=0·9 and at altitudes from sea level to 40 000 ft (12 km) are used in the calculations.

This paper shows the advantages of applying exergy-based analysis and optimisation methods to the synthesis/design and operation of aircraft systems. In particular, an Advanced Aircraft Fighter (AAF) with three subsystems: a Propulsion Subsystem (PS), an Environmental Control Subsystem (ECS), and an Airframe Subsystem – Aerodynamics (AFS-A) is used to illustrate these advantages. Thermodynamic (both energy and exergy based), aerodynamic, geometric, and physical models of the components comprising the subsystems are developed and their interactions defined. Off-design performance is considered as well and is used in the analysis and optimisation of system synthesis/design and operation as the aircraft is flown over an entire mission.

An exergy-based parametric study of the PS and its components is first presented in order to show the type of detailed information on internal system losses which an exergy analysis can provide and an energy analysis by its very nature is unable to provide. This is followed by a series of constrained, system synthesis/design optimisations based on five different objective functions, which define energy-based and exergy-based measures of performance. The former involve minimising the gross takeoff weight or maximising the thrust efficiency while the latter involve minimising the rates of exergy destruction plus the rate of exergy fuel loss (with and without AFS-A losses) or maximising the thermodynamic effectiveness.

A first set of optimisations involving four of the objectives (two energy-based and two exergy-based) are performed with only PS and ECS degrees of freedom. Losses for the AFS-A are not incorporated into the two exergy-based objectives. The results show that as expected all four objectives globally produce the same optimum vehicle. A second set of optimisations is then performed with AFS-A degrees of freedom and again with two energy- and exergy-based objectives. However, this time one of the exergy-based objectives incorporates AFS-A losses directly into the objective. The results are that with this latter objective, a significantly better optimum vehicle is produced. Thus, an exergy-based approach is not only able to pinpoint where the greatest inefficiencies in the system occur but appears at least in this case to produce a superior optimum vehicle as well by accounting for irreversibility losses in subsystems (e.g., the AFS-A) only indirectly tied to fuel usage.

This paper describes a flight parameter based fatigue life analysis approach, which is developed for the Finnish Air Force F-18 fighters. It produces a flight specific fatigue life estimate for structural details using flight parameter data stored by each aircraft. Artificial neural networks are used to model structural response of analyzed details as a function of flight parameters. The analysis development is based on strain gauge data from 25 flights of an instrumented aircraft. The results show a satisfactory accuracy for the fatigue life estimates and prove the concept level analysis capability. The average difference between measured and modelled fatigue life is 21% for the fuselage bulkhead and 30% for the leading edge flap’s hinge area. The total differences in the Finnish Air Force average usage are extremely small, being –2% for the bulkhead and +2% for the leading edge flap.

The paper investigates, by means of a simulation methodology, the flow separating from a 40 degrees backward-swept wing at 9 degrees incidence and Reynolds number of 210,000, based on the wing-root chord length. The Simulation corresponds to LDA, PIV and suction-side-topology measurements for the same geometry, conducted by other investigators specifically to provide validation data. The finest block-structured mesh contains 23·6 million nodes and is organised in 256 blocks to maximise mesh quality and facilitate parallel solution on multi-processor machines. The near-wall layer is resolved, to a thickness of about y+ = 20, by means of parabolised URANS equations that include an algebraic eddy-viscosity model and from which the wall-shear stress is extracted to provide an unsteady boundary condition for the simulation. The numerical solution is in good agreement with the experimental behaviour over the 50-70% inboard portion of the span, but the simulation fails to resolve some complex features close to the wing tip, due to a premature leading-edge vortex breakdown and loss in vortex coherence. The comparisons and their discussion provide useful insight into various physical characteristics of this complex separated wing flow.

One of the challenges of modern engineering design is the amount of data that designers must keep track while performing system analysis and synthesis. This task is particularly important in the design process of complex systems such as novel aerospace systems where Modeling and Simulation play an essential role. The Agile philosophy stems from the field of Software Engineering and describes an approach to development in which requirements and solutions gradually develop through collaboration between self-organising cross-functional teams and end users. Agile Model-Based System Engineering (AMBSE) is the application of the Agile philosophy to Model-Based System Engineering. In this paper, AMBSE is accomplished through the application of the Object-Oriented System Engineering Method (OOSEM). OOSEM employs a top-down scenario-driven process that adopts System Modeling Language (SysML) and leverages the object-oriented paradigm to support the analysis, specification, design, and verification of systems. AMBSE assisted by mathematical modelling and safety assessment techniques is applied to the first design iterations of the main aircraft systems, allowing a comprehensive design exploration. The flight control system was chosen to illustrate the procedure in detail, emphasising the synthesis of a six-degrees-of-freedom model augmented by dynamic inversion control for a hypothetical supersonic transport aircraft satisfying class II MIL-F-8785C handling qualities. It is concluded that AMBSE presents promising properties to support future aircraft development within the current regulatory framework for aircraft design, while enabling a smooth transition from conceptual to preliminary design.

Commercial jets usually have relatively low-aspect-ratio wings, in spite of the associated benefits of increasing the wing aspect-ratio, such as higher lift-to-drag ratios and ranges. This is partially explained by the fact that the wing becomes more flexible by increasing the aspect-ratio that results in higher deflections which can cause aeroelastic instability problems such as flutter. An aeroelastic computational framework capable of evaluating the effects of geometric non-linearities on the aeroelastic performance of high-aspect-ratio wings has been developed and validated using numerical and experimental data. In this work, the aeroelastic performance of a base wing model with 20 m span and 1 m chord is analysed and the effect of changing the wing chord or the taper-ratio is determined. The non-linear static aeroelastic equilibrium solutions are compared in terms of drag polar, root bending moment and natural frequencies, and the change in the flutter speed boundary is assessed as a function of aspect-ratio using a time-marching approach.

An experimental high-aspect ratio wing aeroelastic model with a device to provide a controllable slender body tip mass distribution has been constructed and the model response due to flutter and limit cycle oscillations has been measured in a wind tunnel test. A theoretical model has also been developed and calculations made to correlate with the experimental data. Structural equations of motion based on nonlinear beam theory are combined with the ONERA aerodynamic stall model (an empirical extension of Theodorsen aerodynamic theory that accounts for flow separation). A dynamic perturbation analysis about a nonlinear static equilibrium is used to determine the small perturbation flutter boundary which is compared to the experimentally determined flutter velocity and flutter frequency. Time simulation is used to compute the limit cycle oscillations response when the flutter/LCO control system is ON or OFF. Theory and experiment are in good agreement for predicting the flutter/LCO suppression that can be achieved with the control device.

This paper presents a novel framework, based on traditional system safety modelling approaches, for the representation and comparison of airworthiness aviation frameworks. A disparate array of military airworthiness frameworks have emerged due to a lack of standardisation and the absence of a recognised organisation needed for the harmonisation of military regulatory frameworks.. The complexity and subsequent cost in inter-agency recognition of existing certification programs has led to the establishment of a European forum of military airworthiness authorities. The forum is working towards establishing a common regulatory framework across its European member states. The common framework provides the systematic basis for a certification of military aircraft that can be readily recognised by all of the member states. This will have many cost and efficiency benefits for the EU. The framework and recognition process have recently been accepted as a method for establishing recognition outside of Europe, with some identified shortfalls. This paper establishes a method for overcoming these shortfalls for nations outside of europe. The Product-Behaviour-Process (PBP) Bow-Tie, which is a novel application of the traditional bowtie risk modelling tool, derives test points that capture the airworthiness attestations for the high-level engineering lifecycle processes of design, production and maintenance. The proposed framework is used to provide a comparison between the Australian Defence Force and United States Army regulatory frameworks. The comparative case-study clearly demonstrates the benefit of the PBP Bow-Tie model in its ability to systematically represent the disparate regulatory frameworks. A novel representation of the output is also described, which facilitates a visual comparison of the results. The application of the PBP Bow-Tie framework to the case-study of regulatory frameworks reveals significant differences that need to be addressed in order for inter-agency recognition.

Mitigation of climate change is a challenge to science and society. Here, we establish a methodology, applicable in multi-disciplinary optimisation (MDO) during aircraft pre-design, allowing a minimisation of the aircraft’s potential climate impact. In this first step we consider supersonic aircraft flying at a cruise altitude between 45kfeet (~13·5km, 150hPa) and 67kfeet (~20·5km, 50hPa). The methodology is based on climate functions, which give a relationship between 4 parameters representing an aircraft/engine configuration and an expected impact on global mean near surface temperature as an indicator for the impact on climate via changes in the greenhouse gases carbon dioxide, water vapour, ozone and methane. These input parameters are cruise altitude pressure, fuel consumption, fuel flow and Mach number. The climate functions for water vapour and carbon dioxide are independent from the chosen engine, whereas the climate functions for ozone and methane depend on engine parameters describing the nitrogen oxide emissions. Ten engine configurations are taken into account, which were considered in the framework of the EU-project HISAC. An analysis of the reliability of the climate functions with respect to the simplified climate-chemistry model AirClim and a detailed analysis of the climate functions is given.

The aim morphing of this study is to detect and reconstruct a fault in angle-of-attack sensor and pitot probes that are components in commercial aircrafts, without false alarm and no need for additional measurements. Real flight data collected from a local airline was used to design the relevant system. Correlation analysis was performed to select the data related to the angle-of-attack and airspeed. Fault detection and reconstruction were carried out by using Adaptive Neural Fuzzy Inference System (ANFIS) and Artificial Neural Networks (ANN), which are machine-learning methods. No false alarm was detected when the fault test following the fault modeling was carried out at 0–1 s range by filtering the residual signal. When the fault was detected, fault reconstruction process was initiated so that system output could be achieved according to estimated sensor data. Instead of using the methods based on hardware redundancy, we designed a new system within the scope of this study.

Hypersonic laminar flow past a compression corner has been numerically investigated using time-accurate computational fluid dynamics (CFD) approach. Two flow conditions were considered relevant to high and low enthalpy conditions with a total specific enthalpy of 19MJ/kg and 2·8MJ/kg. The Mach number and unit Reynolds number per metre were 7·5, 9·1 and 3·10 × 105 and 32·2 × 105 respectively. These free stream conditions provided attached, incipiently separated and fully separated flows for ramp angles between θw = 5° to 24°. A grid independence study has been carried out to estimate the sensitivity of heat flux and skin friction in the strong interaction regions of the flow. The investigation was carried out assuming the flow to be laminar throughout and high temperature effects such as thermal and chemical nonequilibrium are studied using Park’s two temperature model with finite rate chemistry. A critical comparison has been made with existing steady state computational and experimental data and the study has highlighted the importance of high temperature effects on the flow separation and reattachment.

This paper proposes a fault-tolerant control (FTC) method based on fast fault observer (FFO) to solve the problem of actuator gain loss fault and stuck fault for hypersonic vehicles. Firstly, an input-output feedback linearisation model is presented that considers parametric uncertainties, control input saturation, disturbances and actuator faults. Secondly, the above factors are defined as an integrated fault item, and an improved fast fault observer is designed to estimate the integrated fault in real time. Finally, the fault-tolerant controller is constructed based on the sliding mode and fault estimation. In case of unknown faults, the effects of gain loss fault or stuck fault happen on elevators and the engine can be quickly processed, Also, the asymptotically stable tracking of the flight output reference command is completed to achieve fault-tolerant control. The final simulation experiment verifies the effectiveness of the proposed method.

The method developed by G. N. Ward for the treatment of slender pointed bodies in a uniform supersonic stream is applied to three special cases.

• (i) Supersonic flow past a body of revolution with thin wings of symmetrical section and of small aspect ratio at zero incidence.

• (ii) Supersonic flow past a body of revolution with plane wings of small aspect ratio set at incidence to the body, the whole being at incidence to the stream.

• (iii) Supersonic flow past a body of revolution with a plane fin of small aspect ratio set at incidence, the whole being at incidence to the stream.

The pressure distribution on the wing has been calculated for a special case of (i) and is given in the Appendix.

There is some evidence to indicate that injuries associated with parachute opening shock are much less frequent than injuries associated with the acceleration loads imposed by ejection seats, even though the stresses presumed to be induced in the human body may often be of the same order. Since all such data is based on operational experience, the details are necessarily vague, and it is hard to draw firm conclusions.

The Aeromedical Research Laboratory, USAF, as part of its studies of aircrew injury, decided to see whether there was any difference between the dynamics of parachute opening when used with a rigid anthropometric dummy of the type normally used to obtain opening shock data, and a human parachutist whose body is naturally more resilient.