The aim of this project was to develop a wind tunnel that enables the study of human performance during various types of sports and physical activities by examining the influence of aerodynamic drag, precipitation, frictional forces and gravitational forces on uphill and downhill travel on a moving substrate. An overall design for a wind tunnel and working section containing a large treadmill was drafted, followed by computational fluid dynamics simulations of flow conditions to assess the design’s feasibility and select from different geometries prior to its construction. The flow conditions in the completed wind tunnel were validated using different flows, speeds and treadmill inclinations. Pilot experiments were carried out using a cross-country skier to investigate the effect of aerodynamic drag on oxygen uptake during double poling and the maximal achieved speed when rolling on a declined treadmill. The purpose was to validate the usefulness of the tunnel. The results showed that flow conditions are acceptable for experiments even in worst-case scenarios with maximal inclined and declined treadmill. Results also showed that aerodynamic drag has a significant impact on the skier’s energy expenditure.

Autonomous contingency management systems, such as a forced-landing system, which reacts appropriately to an engine failure, are important for the safe operation of unmanned aircraft systems. This paper details a method to ascertain the reachability of any possible emergency landing site for a forced landing in steady uniform wind conditions. With knowledge of the aircraft’s state, such as speed, heading, location, and orientation of a landing site, a method to calculate a minimum height loss path is developed based on aircraft glide performance. Wind direction and speed are taken into account using a trochoidal approach by defining the minimum height loss turn path. To facilitate real-time implementation, simplified gliding equations are developed without accuracy loss. The reachability of each site can be calculated as well as how much safety margin an aircraft would have. This method is generic and could also provide decision support for human pilots in forced-landing situations. Two types of aircraft,...

ABSTRACTParticle Image Velocimetry (PIV) has been used to study the complex flowfield created by simulated battle damage to a two-dimensional wing. Computational Fluid Dynamics (CFD) predictions have also been used for validation of internal cavity flow. Two damage cases were selected for the study; both cases were simulated using a single hole with diameters equal to 20% and 40% of the chord, located at the wing half-chord. Wind-tunnel tests were conducted at a Reynolds number of 500,000 over a range of incidences from 0 to 10° with two-component PIV measurements made on three chordwise and three spanwise planes. The PIV data were analysed and compared to CFD data of the same damage cases. The PIV data have shown lower velocity ratios and lower vorticity in the jet compared to past Jet in Cross-Flow experiments and CFD was used to describe the flow features inside the cavity of the wing. It was seen that the wing cavity has large effects on the external flow features, particularly for the 20% damage case. Finally, the flow field data have been related to force balance data. At higher incidence angles, the larger force coefficient increments in both lift and drag can be attributed to the larger wakes and higher jet strengths.

No AccessTechnical NoteSite Selection During Unmanned Aerial System Forced Landings Using Decision-Making Bayesian NetworksMatthew Coombes, Wen-Hua Chen and Peter RenderMatthew CoombesLoughborough University, Loughborough, England LE11 3TU, United Kingdom*Research Associate, Department of Automotive and Aeronautical Engineering; .Search for more papers by this author, Wen-Hua ChenLoughborough University, Loughborough, England LE11 3TU, United Kingdom†Professor in Autonomous Vehicles, Department of Automotive and Aeronautical Engineering; .Search for more papers by this author and Peter RenderLoughborough University, Loughborough, England LE11 3TU, United Kingdom‡Senior Lecturer in Aircraft Aerodynamics, Department of Automotive and Aeronautical Engineering; .Search for more papers by this authorPublished Online:20 Dec 2016https://doi.org/10.2514/1.I010432SectionsRead Now ToolsAdd to favoritesDownload citationTrack citations ShareShare onFacebookTwitterLinked InRedditEmail About References [1] Cox T. 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Moss5 October 2018 | Journal of Air Transportation, Vol. 26, No. 4 What's Popular Volume 13, Number 12December 2016 Metrics CrossmarkInformationCopyright © 2016 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. All requests for copying and permission to reprint should be submitted to CCC at www.copyright.com; employ the ISSN 2327-3097 (online) to initiate your request. See also AIAA Rights and Permissions www.aiaa.org/randp. TopicsAircraft Components and StructureAircraft DesignAircraft OperationsAircraft Operations and TechnologyAircraft SystemsAircraftsArtificial IntelligenceComputing and InformaticsComputing, Information, and CommunicationData ScienceEmergency LandingForced LandingGeographic Information SystemsInformation SystemMachine LearningRobot KinematicsRoboticsTakeoff and Landing KeywordsBayesian NetworksUnmanned Aerial SystemsSafe Forced LandingAirframesLanding Distance AvailableNational Airspace SystemCommercial AircraftPath PlanningCrash LandingCessnaPDF Received26 October 2015Accepted22 September 2016Published online20 December 2016

This paper details a method to ascertain the reachability of known emergency landing sites for any fixed wing aircraft in a forced landing due to engine failure in steady uniform wind conditions. With knowledge of the aircraft's state and parameters, and landing site location and landing direction, the minimum height loss path can be defined. This uses glide performance calculations and a trajectory planner to give a minimum height loss to each landing site. Based on the aircraft's initial altitude it can calculate if the site is reachable, and how reachable it is. The path definition takes into account wind and uses a geometric shape called a trochoid to define the gliding turns in wind. This method is generic enough for use by any aircraft in any wind conditions.

Particle Image Velocimetry (PIV) has been used to map the complex flow field generated by simulated battle damage to a two-dimensional wing. Previous studies have relied on surface flow visualisation techniques to study the flow but here PIV data has enabled the flow field away from the surface to be analysed for the first time. Damage was simulated by a single hole with a diameter equal to 20% of the chord, located at mid-chord. Wind tunnel tests were conducted at a Reynolds number of 500,000 over a range of incidences from 0-10 with two-component PIV measurements made on three span-wise planes; on the damage centre line and o set by 0.5 and 1.0 hole radii. The PIV data was seen to be in good agreement with existing surface flow visualisation showing strong evidence of the formation of a horse shoe vortex, a counter-rotating vortex pair and reverse flow regions. Large variations in the flow structure were observed over the range of incidences tested as the jet transitioned from weak at lower angles to strong at higher angles. The data also revealed some significant differences in the flow compared to classic Jets In Cross-Flow (JICF) behaviour. Notably in the case of battle damage the jet never fully occupies the hole and jet velocity pro le is highly skewed towards the rear of the hole. Additionally, the measured velocity ratios are much less than would be expected for typical JICF. For example, strong jet behaviour is observed at a velocity ratio as low as 0.22 whereas JICF studies would suggest a much higher ratio (> 2) is required. Increasing velocity ratio has been related to a reduction in lift and an increase in drag. At the highest incidence tested (10 ) the velocity ratio of 0.32 resulted in a reduction of the lift coe fficient by 0.18 and an increase in the drag coeffi cient of 0.035.

This paper details a method to ascertain the reachability of known emergency landing sites for any fixed wing aircraft in a forced landing situation. With a knowledge of the aircraft’s state and parameters, as well as a known wind profile, the area of maximum glide range can be calculated using aircraft equations of motion for gliding flight. A landing descent circuit technique used by human pilots carrying out forced landings called high key low key is employed to account for the extra glide distance required for an approach and landing. By combining maximum glide range analysis with the descent circuit, all the reachable landing sites can be determined. X-Plane flight simulator is used to demonstrate and validate the techniques presented.

Wind tunnel tests were carried out on an NASA LS(1)-0417MOD aerofoil with a circular hole simulating gunfire damage. To represent different attack directions, the inclination of the hole axis relative to the aerofoil chord was varied. The hole had a diameter of 20% of the aerofoil chord and its axis was centred at half chord. The aerofoil spanned the wind tunnel to create approximate two-dimensional conditions and balance measurements were carried out at a Reynolds number of 500, 000. Surface flow visualisation and pressure measurements were also carried out. The aerofoil model incorporated a cavity to represent the internal geometry of an aircraft’s wing. Compared to an undamaged wing the addition of damage increased drag, reduced lift and gave a more negative pitching moment. The effects increased with incidence. Adding negative obliquity, where the upper surface hole was moved forward and the lower hole rearwards, increased the magnitude of these effects. Except when combined with extreme negative obliquity, adding skew, where the holes were offset in a spanwise direction, had little measurable effect in terms of aerodynamic coefficients. However, adding skew introduced asymmetry to the flow through the damage.

This paper describes the development of a system for the rapid prototyping of high level control algorithms using an Arduino based commercial off the shelf autopilot called ArduPilot. It is capable of controlling multiple vehicle types, including fixed, and rotary wing aircraft as well as ground vehicles. The inner loop control is performed by ArduPilot, so the high level control can be rapidly prototyped and tested in Simulink, or an embedded system. The ability to conduct tests in software and hardware in the loop has also be developed, to enable safe testing of algorithms, which will speed up the development process. To show its functionality and ability to assist with the development process of algorithms, ArduPilot is used with a remote controlled aircraft in simulation and in real world testing to verify newly developed high level algorithms for UAVs.

Numerical studies have been carried out on a battle damaged NACA 641-412 half wing of aspect ratio 8.2 at a Reynolds number of 5.5 x 105. The simulated gunfire damage was represented by a single hole with a diameter of 0.2 wing chord. The hole was centred at half chord and at spanwise locations of 450mm and 650mm from wing root. Computational Fluid Dynamics (CFD) results have been compared with experimental data and a general overall good agreement has been obtained, in the flow features introduced by the damage. Furthermore, the flow field inside the damage hole and in the region downstream of the jet have been analysed in more detail using CFD.

The importance of energy and specific excess power in assessing the maneuverability of combat aircraft is discussed. This is followed by a discussion on climb and turn performance, which extends the methods presented in other sections. In particular, high angle climbs and the climbing turn are considered. The use of performance metrics to assess aircraft performance is described, and both traditional (e.g., specific excess power plots) and new metrics (e.g., combat cycle time) are described. The article concludes by considering features which have been adopted on the latest combat aircraft, namely relaxed stability, the canard configuration and thrust vectoring.

The half model wind tunnel technique suffers from aerodynamic loses due to the interaction of the tunnel wall boundary layer with the flow over the model and the formation of a horseshoe vortex in the model-floor junction. The vortex is believed to contribute to the degradation of the half model aerodynamic performance. An attempt was made to reduce the aerodynamics losses by modifying the junction horseshoe vortex through the use of localized suction just upstream of the model leading edge. Wind tunnel tests on a rectangular and untwisted, wing only, modified LS(1)-0413 half model were conducted at Reynolds numbers of 0.44 x 10(6), 0.88 x 10(6) and 1 x 10(6). Without suction the force and moment balance measurements of the half model showed the anticipated deviation from full model values, e.g. lower lift curve slope and higher drag values. Effects of localized suction were limited to Reynolds number of 0.44 x 10(6) and improvements were seen only near stall angles of attack. Flow visualization for the no suction case showed that a horseshoe vortex did not exist over much of the incidence range for this particular model and hence there was little room for suction to effect junction flow. Near stall, suction removed the horseshoe vortex around the upper surface of the model and significantly reduced flow separations occurring in the model-floor junction, leading to the improved stall characteristics.

Wind-tunnel tests have been carried out on a battle-damaged NACA 64(1)-412 half-wing aspect ratio of 8.2. The simulated gunfire damage had a diameter of 0.2 wing chord and was located at midchord and at one of two spanwise locations. Tests were carried out at a Reynolds number of 5.5 x 10(5). Compared with an undamaged wing, the damage resulted in reduced lift, increased drag and a positive increase in pitching moment at zero lift. Moving the damage to near the tip reduced the magnitude of these effects. Using the static pressure difference between the upper and lower surfaces of the undamaged wing allowed the data from the present study to be successfully compared with previously published drag and lift data for a two-dimensional damaged airfoil. Tests on wings with aspect ratios of 6.2 and 10.3 produced similar trends in the aerodynamic characteristics and showed that the use of static pressure difference was equally effective in allowing comparisons with two-dimensional data.

Wind tunnel tests are reported that investigate three aspects of aerodynamic flows through battle damaged airfoils. The first aspect investigated was the effect of camber. This showed that reducing camber weakened the strength of the jet flow through the damage and delayed the onset of strong jet flows to higher incidences. The second investigation used five hole probe measurements to survey the flow field on a battle damaged flat plate airfoil. The measurements indicated that the use of the jet-to-freestream velocity ratio is a poor criteria for determining whether damage flows have undergone transition to strong jets. Finally, the influence of a star shaped hole to simulate more realistic battle damage was investigated. It was shown that in terms of damage flow characteristics and changes in lift, drag, and pitching moment coefficients, the use of a circular hole is a reasonable simulation of battle damage.

Although test rig data exists for ¼ circle turning vanes, the actual performance of these vanes once installed in a wind tunnel, and the extent to which test rig results are replicated, is rarely known. This paper compares pressure loss coefficient and velocity profile data from a vane test module with measurements taken in the low speed wind tunnel described in Ref. 1.

Two dimensional wind tunnel tests were conducted on NACA aerofoils with simulated battle damage. The first set of tests showed that positive deflection of a trailing edge flap increased the strength of the flow through the damage For a given incidence, increasing flap deflection increased the drag coefficient relative to the undamaged aerofoil and reduced the lift coefficient. The second set of tests demonstrated that similar effects can be obtained by increasing camber. The final set of tests investigated the effectiveness of battle damage repair schemes. Realistic schemes, incorporating both upper and lower surfaces repairs, were shown to achieve reductions in drag coefficient and increases in lift coefficient, which in both cases were significant. Repairs to only one surface (i,e, upper or lower) were also shown to make significant aerodynamic improvements, which were attributed to the lack of flow through the damage.

The flow on a full span NACA 641-412 airfoil with right triangle, inverse triangle and star shaped damage was experimentally investigated using balance measurements and flow visualisation. Generally, when compared with an undamaged model, increasing incidence for a damaged model resulted in increased loss of lift coefficient, increased drag coefficient and more negative pitching moment coefficient. The results are compared to each other and also to the results of circle shaped damage from previous work. The experiments showed that for all damage cases the flows could be categorised as weak, transitional or strong jets, with the main features of these flows identical to circle shaped damage. Adding multiple sharp edges (star) had only limited influence on the observed flow characteristics. For all damage cases the jet exited from the rear of the damage and its size was determined by the width of the rear part of the hole.

AbstractThe design methodology and performance of Loughborough University’s new 1·9m × 1·3m, indraft wind tunnel is discussed in the following paper. To overcome severe spatial and financial constraints, a novel configuration was employed, with the inlet and exit placed adjacent to each other and opened to atmosphere. Using a fine filter mesh, honeycomb, two turbulence reduction screens and a contraction ratio of 7·3, flow uniformity in the working area of the jet at 40ms-1is shown to be within 0·3% deviation from the mean velocity, with turbulence intensity in the region of 0·15%. Working section boundary layer characteristics are shown to be consistent with that of a turbulent boundary layer growing along a flat plate, which originates at the point of inflection of the contraction. A maximum velocity of 46ms-1was achieved from a 140kW motor, compared to a prediction of 44ms-1, giving an energy ratio of 1·42. Comparison between theoretical and measured performance metrics indicate differences between the way modules perform when part of a wind tunnel system compared to data gathered from test rigs.