This paper presents a new adaptive algorithm for the statistical filtering of miniaturised inertial sensor noise. The algorithm uses the minimum variance method to perform a best estimate calculation of the accelerations or angular speeds on each of the three axes of an Inertial Measurement Unit (IMU) by using the information from some accelerometers and gyros arrays placed along the IMU axes. Also, the proposed algorithm allows the reduction of both components of the sensors’ noise (long term and short term) by using redundant linear configurations for the sensors dispositions.

A numerical simulation is performed to illustrate how the algorithm works, using an accelerometer sensor model and a four-sensor array (unbiased and with different noise densities). Three cases of ideal input acceleration are considered: 1) a null signal; 2) a step signal with a no-null time step; and 3) a low frequency sinusoidal signal.

To experimentally validate the proposed algorithm, some bench tests are performed. In this way, two sensors configurations are used: 1) one accelerometers array with four miniaturised sensors (n = 4); and 2) one accelerometers array with nine miniaturised sensors (n = 9). Each of the two configurations are tested for three cases of input accelerations: 0ms−1, 9·80655m/s2 and 9·80655m/s2.

Piezoelectric wafer active sensors (PWAS) are lightweight and inexpensive transducers that enable a large class of structural health monitoring (SHM) applications such as: (a) embedded guided wave ultrasonics, i.e., pitch-catch, pulse-echo, phased arrays; (b) high-frequency modal sensing, i.e., electro-mechanical impedance method; and (c) passive detection. The focus of this paper is on the challenges posed by using PWAS transducers in the composite laminate structures as different from the metallic structures on which this methodology was initially developed. After a brief introduction, the paper reviews the PWAS-based SHM principles. It follows with a discussion of guided wave propagation in composites and PWAS tuning effects. Then, the mechanical effect is discussed on the integration of piezoelectric wafer inside the laminate using a compression after impact. Experiments were performed on a glass fibre laminate, employing PWAS to measure the attenuation coefficient. Finally, the paper presents some experimental and multi-physics finite element method (MP-FEM) results on guided wave propagation in composite laminate specimens.

Aerodynamic design of an innovative supersonic three-dimensional air-intake including a starting device has been carried out. The intake concept is based on the use of an external compression ramp in the form of a caret waverider. In a design flow regime, this ramp generates a plane oblique shock wave lying on its swept leading-edges and a subsequent two-dimensional isentropic compression wave. The intake starting is provided with a device special flaps of which in the throat are opened slightly in order to enlarge it and to form here both longitudinal and cross slots for air bleed and boundary-layer diversion. The same device can also serve for regulation of intake operating regimes depending on the flight velocity. The experimental model intake with a design Mach number MD = 2 has been developed and tested at flow Mach numbers 1·5, 1·75, and 2·0 in the T-313 supersonic wind tunnel based at ITAM. Flow patterns about the intake have been observed, and flow parameters characterising the intake performance have been measured. The experimental results obtained have confirmed the workability of the starting/regulating device; the data on the efficiency of the model intake have demonstrated its good characteristics in terms of the total pressure recovery.

Flight simulators are integral to the design/development, testing/qualification, training and research communities and their utilisation is ever expanding. The use of flight simulation to provide a safe environment for pilot training, and in research and development, must be underpinned by quantification of simulator fidelity. While regulatory simulator standards exist for flight training simulators and new standards are in development, previous research has shown that current standards do not provide a fully quantitative approach for assessing simulation fidelity, especially in a research environment. This paper reports on progress made in a research project at the University of Liverpool (Lifting Standards), in which new predicted and perceptual measures of simulator fidelity have been developed. The new metrics have been derived from handling qualities engineering practice. Results from flight tests on the National Research Council (Canada) Bell 412 ASRA research aircraft and piloted simulation trials using the HELIFLIGHT-R simulator at Liverpool are presented to show the efficacy of adopting a handling qualities approach for fidelity assessment. Analysis of the new metrics has shown an appropriate degree of sensitivity to differences between flight and simulation.

The use of machine vision to guide robotic spacecraft is being considered for a wide range of missions, such as planetary approach and landing, asteroid and small body sampling operations and in-orbit rendezvous and docking. Numerical simulation plays an essential role in the development and testing of such systems, which in the context of vision-guidance means that realistic sequences of navigation images are required, together with knowledge of the ground-truth camera motion. Computer generated imagery (CGI) offers a variety of benefits over real images, such as availability, cost, flexibility and knowledge of the ground truth camera motion to high precision. However, standard CGI methods developed for terrestrial applications lack the realism, fidelity and performance required for engineering simulations.

In this paper, we present the results of our ongoing work to develop a suitable CGI-based test environment for spacecraft vision guidance systems. We focus on the various issues involved with image simulation, including the selection of standard CGI techniques and the adaptations required for use in space applications. We also describe our approach to integration with high-fidelity end-to-end mission simulators, and summarise a variety of European Space Agency research and development projects that used our test environment.

This paper presents an experimental investigation of the evolution of the leading-edge vortex and spanwise flow generated by an insect-like flapping-wing at a Reynolds number relevant to flapping-wing micro air vehicles (FMAVs) (Re = ~15,000). Experiments were accomplished with a first-of-its-kind flapping-wing apparatus. Dense pseudo-volumetric particle image velocimetry (PIV) measurements from 18% – 117% span were taken at 12 azimuthal positions throughout a flapping half cycle. Results revealed the formation of a primary leading-edge vortex (LEV) which saw an increase in size and spanwise flow (towards the tip) through its core as the wing swept from rest to the mid-stroke position where signs of vortex breakdown were observed. Beyond mid-stroke, spanwise flow decreased and the tip vortex grew in size and exhibited a reversal in its axial direction. At the end of the flapping half cycle, the primary LEV was still present over the wing surface, suggesting that the LEV remains attached to the wing throughout the entire flapping half cycle.

Airbus and Boeing are cooperatively presenting this topic dealing with transport aircraft stalls. The paper will begin by defining a stall, followed by a review of requirements, predictive validation and flight testing. There are various ways of designing modern jet transports for the stall regime such as aerodynamic approaches, flight deck indications, and augmentation control laws to deal with the high angle-of-attack (α) arena. The goal of augmented control laws for high α is common – no full aerodynamic stall or loss of climb performance should occur in the operational flight envelope, in Normal flight control modes. The validation techniques employed in preparation for a flight test campaign will follow. These include flight characteristic predictions based on wind-tunnel data as well as pilot-in-the-loop simulation rehearsals. The preparation for flight testing will be reviewed from both the engineer and pilot viewpoints. This will be followed by a review of various flight testing that has been conducted. The paper will close with a brief foray into what the future of transport stalls could be – perhaps protection features in degraded flight control modes? What are the benefits as well as drawbacks to increased augmentation for high α?

Weight estimation methods are categorised in different classes based on their level of fidelity. The lower class methods are based on statistical data, while higher class methods use physics based calculations. Statistical weight estimation methods are usually utilised in early design stages when the knowledge of designers about the new aircraft is limited. Higher class methods are applied in later design steps when the design is mature enough. Lower class methods are sometimes preferred in later design stages, even though the designers have enough knowledge about the design to use higher class methods. In high level multidisciplinary design optimisation (MDO) fidelity is often sacrificed to obtain models with shorter computation times. There is always a compromise required to select the proper weight estimation method for an MDO project.

An investigation has been performed to study the effect of using different weight estimation methods, with low and medium levels of fidelity, on the results of a wing design using multidisciplinary design optimisation techniques. An MDO problem was formulated to design the wing planform of a typical turboprop and a turbofan passenger aircraft. The aircraft maximum take-off weight was selected as the objective function. A quasi-three-dimensional aerodynamic solver was developed to calculate the wing aerodynamic characteristics. Five various statistical methods and a quasi-analytical method are used to estimate the wing structural weight. These methods are compared to each other by analysing their accuracy and sensitivity to different design variables. The results of the optimisations showed that the optimum wing shape is affected by the method used to estimate the wing weight. Using different weight estimation methods also strongly affects the optimisation convergence history and computational time.

An experimental study incorporating the use of the Background-Oriented Schlieren (BOS) technique was performed to measure the density field of a rectangular supersonic jet. This technique is easier to set up than conventional schlieren since the optical alignment involving the various mirrors, lenses and knife-edge is replaced by a background pattern and a single digital camera. The acquired images which contain information of density gradients in the flow are solved as a Poisson equation and further processed using deconvolution and tomographic algorithms to generate a 3D domain which contains information about the actual density. 2D slices can then be extracted to quantitatively visualise the density along any required planes. The results from supersonic axisymmetric jets are used for validation of the code; these show excellent agreement with pre-validated CFD data. The results for a rectangular supersonic jet are then obtained. These show good agreement with the CFD data, in terms of shock-cell spacing and overall structure of the jet. The technique has proved useful for investigating axis-switching, a phenomenon generally associated with non-axisymmetric jets.

At the heart of a flight simulator resides the mathematical representation of aircraft behaviour in response to control inputs, atmospheric disturbances and system inputs including failures and malfunctions. While this mathematical model can never be wholly accurate, its fidelity, in comparison with real world behaviour, underpins the usefulness of the flight simulator. The present paper examines the state of the art achieved in validating mathematical models for helicopter simulators, addressing the strengths and weaknesses of the present European standard for the qualification of helicopter flight simulators, JAR FSTD-H (previously JAR-STD-1H/2H/3H). Essential questions are examined, such as: What is the required model fidelity to guarantee a simulation is sufficiently representative to be fit for purpose? Are the tolerances set in the current standards fine enough that they lead to only minor changes in handling qualities? What is an acceptable tuning process for the simulation? What is the effect of modelling fidelity on the overall pilot control strategy? What is the relationship between the settings of the simulator cueing environment and the behaviour of the pilot? What is the industrial experience on qualification of flight simulators that might usefully inform developments? Many of these questions were addressed in Europe in a previous GARTEUR Action Group (AG) HC/AG-12 the results of which are documented in this paper. Solutions are proposed for improving the current JAR-FSTD standard with respect to validation of mathematical models.

This paper will address the engineering performance differentiators for an F1 car and highlight the difference aerodynamics can make to that performance. It will also consider some basic aerodynamic challenges and the main tools used for aerodynamic exploration by teams.

This paper assesses the application of a new method for system stability analysis, the weight functions method, to the longitudinal and lateral motions of a High Incidence Research Aircraft Model. The method consists of finding the number of weight functions that is equal to the number of differential equations required for system modelling. The aircraft’s stability is determined from the sign of the total weight function; which should be negative for a stable model. The Aero-Data Model In Research Environment (ADMIRE) simulation, developed by the Swedish Defence Research Agency, was used for the aerodynamic aircraft modelling, with the following configurations: Mach number = 0·25, altitude = 500m, angle-of-attack [–10 to 30]°, elevon deflection angle [–30 to 30]°, canard deflection [0° and 25°] and rudder deflection angles [–30° and 30°]. These flight configurations were selected because they are among the flight conditions for Cat. II Pilot Induced Oscillation (PIO) criteria validation, performed on the FOI aircraft model presented in the PIO Handbook by the Group for Aeronautical Research and Technology in Europe, Flight Mechanics/Action Group 12. This aircraft model has a known instability for longitudinal and lateral motions and so a control law was introduced to stabilise its flight.

An optimisation procedure coupled with computational fluid dynamics (CFD) is proposed to minimise the aerodynamic drag and to improve the static and dynamic stabilities of generic rounds at supersonic speeds (Mach 1·5 to 4). First, the Active-set algorithm, Sequential Quadratic Programming (SQP) is used as the optimisation method for drag minimisation. The objective function is the zero-lift drag computed from a semi-empirical solution. The constraints are based on the geometric restrictions of the body. CFD is then employed to validate the accuracy of the drag prediction from the semi-empirical solution and to incorporate the stability requirements into the optimisation process. A supersonic round body is considered as an example application. The optimised body provides up to 15% drag reduction and 46% increase in gyroscopic stability while remaining dynamically stable over the whole range of the operating Mach numbers.

Air-to-ground strike has become one of main forms of modern warfare, and the guided bomb release is an important and key part of the attack. The guided bomb release planning aims to accomplish a precise target hit in enemy’s region while guarantee the pilot’s safety. We propose a robust Monte Carlo method by taking error perturbations into consideration, comparing to the traditional sequential quadratic programming method under extreme conditions. At the same time using a distributed virtual physics environment we can obtain much more detailed realism relative to the conventional simulator running on a single machine. The experimental results verify that Monte Carlo methods can improve hit probability and weapon efficiency significantly. Furthermore, the 3D visualised environment plays a very important role in training pilots, so this simulator will decrease the cost and time requirements of physical experiment that are not always compatible with strict military task.

The current flight management system (FMS), CMA-9000, from CMC Electronics-Esterline, only optimises the vertical flight profile in terms of the speed of the aircraft. This article defines a methodology that optimises the speeds and altitudes for the vertical profile, obtaining a trajectory that reduces the global flight cost.

The performance database (PDB) provided by CMC Electronics-Esterline is presently used on aircraft such as the Lockheed L-1011, the Airbus A310 and the Sukhoi Superjet 100 Russian regional jet. The PDB is used as the reference to design different trajectory optimisation algorithms to obtain the altitude where the aircraft fuel efficiency is the best. These algorithms are compared with the part-task trainer (PTT), simulator that represents the FMS CMA-9000, supplied by CMC Electronics-Esterline as well.

To validate the results, the FlightSIM® software is used, which considers a complete aircraft aerodynamic model for its simulations, giving accurate results and very close to reality.

A scale model of aerial refuelling drogue is built and tested to investigate active drogue stabilisation and control concepts. A set of aerodynamic control surfaces is implemented to a conventional, aerodynamically stabilised refuelling drogue. The control surfaces are designed to reduce the response of the drogue motion to atmospheric disturbance and tanker motion. This paper presents the details of design concept and experimental results based on wind-tunnel testing of a ⅓ scale model fabricated for this study. To investigate the proposed active control concept, a dynamic test rig is built for wind-tunnel experiment. The rig basically represents a hose-drogue system in terms of a 4 degree-of-freedom (DoF), single link pendulum model. System identification technique is used to obtain a drogue dynamic model, based on which a feedback control law is developed. Closed-loop dynamic testing is carried out to evaluate the effectiveness of the aerodynamic surface control module and feedback control law.

During the decay process the shape of wake vortices changes significantly which has an influence on the encounter characteristics, hence on the encounter hazard. In order to evaluate the influence of vortex deformation on the wake encounter hazard, in-flight simulations with the DLR research aircraft ATTAS were carried out. For a realistic analysis of wake encounters flow fields of matured vortices were generated with large eddy simulations. These flow fields were used for the determination of histories of induced forces and moments acting on the wake encountering aircraft. The force and moment histories were then fed into the equations of motions of the non-linear six degree-of-freedom in-flight simulation of the DLR research aircraft ATTAS. In order to compare different stages of vortex deformation, encounters were simulated in flight with wavy vortices and vortex rings. The most important benefit of the in-flight-simulation is the realistic environment, which enables a realistic assessment of pilots’ encounter acceptance. The analysed scenario was of a wake encounter during final approach. The encounter conditions correspond to separation distances of about 4nm and 5nm behind an aircraft of the ‘heavy’ category. During the encounters the ATTAS was flown under manual control. Altogether 31 encounters were simulated in flight, 9 with wavy vortices and 22 with vortex rings.

As Large Eddy Simulation is increasingly applied to flows containing complex geometry, grid generation becomes difficult and time consuming when using software originally developed for RANS flow solvers. The traditional ‘pipeline’ approach of grid generation → solve → visualise entails the time consuming transfer of large files and conversion of file formats. This work demonstrates a grid generation methodology developed specifically to be integrated with parallel LES. The current approach is to use a Cartesian grid with adaptive refinement based upon geometry intersection, surface detail and surface curvature. The grid is defined by an octree data structure with the geometry defined by triangular facets using the STL file format. The result is a set of ‘cubical’ subdomains, each with identical numbers of cells and uniform distributions within the cube. Some subdomains will be entirely fluid and can be solved using straightforward CFD techniques, whilst some cubes will be cut by the surfaces. Individual cells are then tagged as ‘solid’, ‘fluid’ or ‘cut’ with the solver expected to use an immersed boundary approach to model the surface. A key feature is the design of the algorithm to be parallelisable on both shared and distributed memory systems. The distributed memory parallel dynamically partitions the grid as it is being generated, so that the partitioning is suitable for a subsequent flow solver. Grid generation testing has been carried out on a variety of input CAD files ranging up to 350,000 facets. A landing gear case shows how the grid generator correctly finds the fluid inside of the tire and other cavities within the hub. In scalar mode, a grid with 4,916 cubes and 468 million cells is generated in less than 100 seconds, whilst in parallel on 32 processor cores this can be achieved in 4·6 seconds.

Boundary-layer noise produced by rough surfaces is a potential contributor to airframe noise. In this paper, an attempt is made to assess the surface roughness noise for a conceptual Silent Aircraft design SAX-40 using a prediction model described in previous theoretical work (Liu and Dowling(12)). Estimates of three idealised test cases show that surface roughness could produce a relatively high noise level and enhance the trailing-edge noise somewhat. Roughness height and roughness density are two significant parameters which affect surface roughness noise, with roughness height having the more dominant effect. The distribution of roughness noise sources decreases in the streamwise direction on the SAX-40 surface. Two candidate rough surfaces are selected for SAX-40 to keep surface roughness noise at a negligible level and to meet an aggressive noise target.