Within the framework of a major European co-operative research project on rotorcraft aerodynamics and acoustics (Helishape), a parametric model rotor test was conducted in the open test section of the DNW using DLR's MWM test rig, and a highly instrumented model of a fully articulated ECF rotor equipped with blades of advanced design with two exchangeable blade tips. One set of blade tips (7A) was rectangular, while the other (7AD1) was a swept-back parabolic-anhedral shape. The objectives of this experimental research were to evaluate noise reduction techniques (conceptionally by variation of rotor speed, dedicated tip shapes and advanced aerofoils, as well as operationally, by identifying low noise, BVI-minimising, descent procedures) and to validate the partners’ aerodynamic and aeroacoustic codes. A comprehensive set of simultaneous acoustic and aerodynamic blade surface pressure data and blade dynamic and performance data was measured. In addition, valuable information on the tip-vortex geometry and blade-vortex miss distance was obtained by LLS flow visualisation. The experimental equipment, the test procedures, and the test matrix are briefly described. A survey of the main results is presented and the trends of the most important parameter variations for both rotors are discussed.

Events in the aerodynamic development of the multi-service V-22 Osprey tiltrotor are reviewed, with emphasis on recent flight test findings and the solutions developed. A discussion of the performance in the USMC, Navy and SOCOM missions is presented.

This paper describes the development of a coupled rotor-flexible fuselage model which is suitable for simulating vibration reduction based on the active control of structural response (ACSR) approach. The rotor is an Nb-bladed aeroelastic model, with coupled flap-lag-torsional dynamics for each blade. Moderate blade deflections are included, together with complete coupling between rotor and fuselage dynamics. This aeroelastic response model is combined with a control algorithm based on an internal model principle. The control scheme effectively reduces vibrations to 0·05g levels or lower, using reasonable actuator forces. The baseline vibration levels in the fuselage are relatively high. This is due to the lack of damping modelling in the fuselage. With the actuators engaged, the hub loads remain virtually unchanged and therefore this control approach has no influence on vehicle airworthiness.

This paper describes a method for using inverse simulation to obtain a preliminary assessment of helicopter handling qualities. Formal descriptions of standard manoeuvres, defined to establish the handling qualities of military helicopters, are used to drive an inverse simulation of a subject helicopter. The simulation generates the controls and states of the helicopter as it executes the manoeuvre and the results may be used to calculate values of quickness, a parameter defined to measure responsiveness. Initial results reveal that in the context of inverse simulation quickness is independent of vehicle configuration when, as specified in the requirements, the quickness is based on the helicopter's kinematic states. An alternative quickness parameter, associated with the control displacements required to fly the manoeuvre is shown to be capable of discriminating between the pilot workload involved in flying two different configurations through the same manoeuvre.

The aerodynamic interference between oscillating models and the walls of slotted windtunnels is controlled by the homogeneous wall boundary condition, which has hitherto been uncertain. In this report a new homogeneous boundary condition for slotted tunnels is suggested which incorporates the important effects of plenum chamber depth and slot parameter.

In particular, this new boundary condition controls the transverse resonance frequencies of the tunnel, which represent a severe form of wall interference. Hence a comparison of measured resonance frequencies with predictions based on such a boundary condition provides a good test of its correctness.

The new homogeneous boundary condition complements that obtained previously for a perforated tunnel and is essentially similar in character. The boundary condition is applicable to oscillatory flows at frequencies ranging from quasi-steady values to multiples of the first resonance frequency.

Since the 1970s, large passenger aircraft design has evolved by modest, but commercially significant, incremental change following the underwing pod concept pioneered by the Boeing 707. In the 21st century, the market is likely to require greater passenger capacity to deal with air-side congestion and higher performance and operational efficiency to justify investment and conserve finite fossil fuel resources. Public opinion will require far greater emphasis on control of noise and engine exhaust pollution. There is likely to be much increased emphasis on pollution in the upper-atmosphere and its environmental impact. In such a situation, it is questionable whether the evolutionary design route can produce the necessary advances and this must stimulate the search for radical design alternatives.

A revolutionary approach, involving the delta planform combined with wing laminar flow control and its impact on overall design, is given a preliminary study in this paper. This has been carried out by assuming that major improvements in drag can be obtained by extensive laminarisation. Using non-dimensional methods, the resulting broad interactive impacts on airframe and engine design and performance are derived. The effect of assuming varying quantities of low-energy air sucked from the foot of the boundary layer is studied and suction-system performance examined.

In addition to the large potential improvements in range, a strong relationship between lower drag, lower cruising altitude and lower cruise engine size is identified. In the study, several factors emerge which may combine to drive large aircraft design towards low aspect ratio and the integrated delta wing planform. This paper is intended as a stimulus and a basis for further study and research.

A method for the calculation of steady or unsteady compressible flows is presented. The procedure, based on a cartesian cut cell approach and multi-dimensional high resolution upwind finite volume scheme, can cope with static or moving body problems having arbitrarily complex geometries. The method is described in two parts. In Part A, we discuss the cartesian cut cell approach and upwind finite volume scheme for static body problems. The method is validated on test problems involving both steady and unsteady compressible flows and then applied to some practical problems. The extension of the method to moving body problems is presented in Part B (pp 57-65).

The presence of numerical noise in engineering design optimisation problems inhibits the use of many gradient-based optimisation methods. This numerical noise may result in the inaccurate calculation of gradients which in turn slows or prevents convergence during optimisation, or it may promote convergence to spurious local optima. The problems created by numerical noise are particularly acute in aircraft design applications where a single aerodynamic or structural analysis of a realistic aircraft configuration may require tens of CPU hours on a supercomputer. The computational expense of the analyses coupled with the convergence difficulties created by numerical noise are significant obstacles to performing aircraft multidisciplinary design optimisation. To address these issues, a procedure has been developed to create noise-free algebraic models of subsonic and supersonic aerodynamic performance quantities, for use in the optimisation of high-speed civil transport (HSCT) aircraft configurations. This procedure employs methods from statistical design of experiments theory and response surface modelling to create the noise-free algebraic models. Results from a sample HSCT design problem involving ten variables are presented to demonstrate the utility of this method.

For a long time there has been growing awareness in the international aviation community of the considerable danger which an encounter with the atmospheric phenomenon of windshear can bring to an aircraft in flight. Considerable statistical evidence is now available which points to the fact that aircraft windshear encounters in initial climb, final approach, or landing are extremely hazardous, often leading to fatal accidents. The paper presents details of a preliminary design for an airborne windshear detection system suitable for use in general aviation aircraft. First, an elementary explanation of windshear, and its most dangerous form, the microburst, is given, together with a short account of the hazards that such atmospheric phenomena can present to aircraft in flight, particularly at takeoff and landing. Then a novel windshear detection algorithm is described and associated simulation results are presented. The algorithm is based upon observer theory and uses only a restricted number of measurements. The system is shown to provide very good estimates of the horizontal and vertical components of some windshear encounters. These estimates of the windshear components are then used to provide the pilot with a warning of the presence of windshear together with an indication of its severity. Digital simulation has been used to show the effectiveness of the proposed design.

The contemporary approach to helicopter mathematical modelling in simulation for handling qualities applications is to represent each blade individually. This allows incorporation of effects not possible with a multiblade disc representation of the rotor system. This paper addresses the validation of such a model, which is uniquely performed for large amplitude manoeuvres, using a recently-developed approach to inverse simulation. The method is reviewed in the context of model validation, and comparisons between simulation and actual data are presented for a Puma helicopter executing sidestep manoeuvres in low-speed flight. The focus for interpreting the results is in the area of wake modelling. It is concluded that further developments are necessary in this area if helicopter flight mechanics models are to be used in simulation of role-related flight.

The flow structure on the upper side of a delta wing is extremely complex. At moderate angle of attack the flowfield is dominated by organised vortical flow structures emanating from the leading edge. The pressure distribution created on the wing surface by these leading edge vortices causes an increment of lift which may be a relevant percentage of the total wing lift, depending on sweep angle.

Delta wing performance is conditioned, however, by a phenomenon known as vortex breakdown or vortex bursting which appears at high angle of attack. This leads to a drastic change in the flowfield which influences the trend of the aerodynamic coefficients.

With the aim of giving a contribution to the understanding of the phenomenon of vortex breakdown, a 65° delta wing has been extensively tested in the low speed windtunnel of Politecnico di Torino (TPI) both in static and dynamic conditions, under forced oscillatory motions.

Many experiments have been carried out in an effort to understand the factors which can affect the breakdown by varying angle of attack, sideslip angle and Reynolds number.

Flow visualizations have been performed using helium bubble tracers. This technique showed a very good capability for visualizing vortical flows and breakdown, both in static and dynamic conditions.

For the static case, pressure distributions are presented, correlated to force measurements and flow visualizations, at different angles of attack and sideslip for different flow conditions.

For the dynamic case force measurements are shown, compared with flow visualizations, for different angles of attack and for different oscillation frequencies.

In 1995, 19 western built jet aircraft were totally destroyed in air crashes which killed 383 passengers and 39 crew members. Although this rate of loss has been steady for the past 10 years, there were three major losses in 1996 — the famous Valujet and TWA aircraft in the United States and the world's worst midair collision in history New Delhi, India, where a Saudia Boeing 747 with 312 persons aboard collided with a Kazak aircraft carrying 37 passengers and crew. All on board were killed.

This paper examines the feasibility of using neural networks to represent the effects of human operators in computer models of complex man-machine systems. In the suggested approach, data from man-in-the-loop simulators are used to train the networks. The method has been tested on several typical data sets using a stand-alone prototype system, consisting of a three-layer feed-forward network with a Chemotaxis training algorithm. Successful results have been obtained, and these can be used to place constraints on the quality, quantity and type of simulator data required in future applications.

Windtunnels are generally used for testing models of solid, aerodynamic objects, such as aircraft. The test articles are generally manufactured to have a representative external profile but provide sufficient strength to survive the harsh windtunnel environment without damage to the test article or tunnel. It is somewhat unusual, however, to test lightweight, high drag, fabric models (i.e. parachutes) in supersonic flow.

This paper describes the testing carried out on the parachutes for the Huygens probe, the techniques developed for the testing and the lessons learned.

Tests were carried out to characterise the drag coefficient of the parachute designs throughout the Mach range of 0.05 to 1.5, and the stability coefficients at incidences of up to 15° at low Mach numbers.

An implicit cell-vertex upwind method is presented for the solution of the two-dimensional Euler equations for supersonic flows. It is based on the two-point difference scheme introduced by S.F. Wornom, but contains many significant developments. Particular attention is given to the implementation of boundary conditions, to ensure that no “numerical” boundary conditions are required, and “image” cells are avoided, making the scheme ideal for use in conjunction with a multiblock approach. Examples are given for quasi-one-dimensional nozzle flows, and in two dimensions, oblique shock reflection, a shock interaction problem, and a supersonic intake problem. The accuracy and efficiency of the scheme is demonstrated, along with its ease of use.

The computation of slowly moving Shockwaves is considered. A numerical error is introduced into the solution during an explicit computation, which results in the generation of spurious entropy. The generation of this spurious entropy and its effect on the pressure wave in Euler flows is examined, in one dimension by considering a moving shock in a constant area duct, and in two dimensions by considering an unsteady aerofoil flow. An efficient temporal discretisation is presented, whose advantages over the explicit scheme are two-fold. First, the numerical entropy generation due to shock motion is eliminated, and second, the CPU requirements for a two-dimensional unsteady computation are reduced by a factor of five from that required for the explicit scheme, for no loss of accuracy.

Closed form solutions for low thrust orbit raising are well known for mission analysis and vehicle performance prediction. This paper provides extended solutions which include nodal regression through the effect of the J2 oblateness term in the geopotential. Since regression is a strong function of orbit radius, the resulting change in ascending node angle is coupled to the in-plane orbital motion. However, it is demonstrated that the resulting dynamical equations may be separated and solved sequentially. Nodal regression during orbit raising is of importance for solar electric vehicles due to the resulting variation of sun aspect angle and eclipse conditions.

Low speed windtunnel measurements were made on a 50° delta wing with leading edge vortex flaps. Improvements in the lift/drag ratio were obtained by deflecting the leading edge vortex flap. Comparisons were made between the previously measured 60° and 70° delta wing results and the present 50° wing results. Improvements in the lift/drag ratio of the 50° delta wing were attained over a wider lift coefficient range than for the 70° delta wing. The highest lift/drag ratio for the 50° delta wing is achieved when the flow attaches to the flap surface without any large area of separation. Estimations of the aerodynamic forces were also made using a quasi-vortex lattice method coupled with the leading edge suction analogy for the 50°, 60° and 70° delta wings. The results obtained from this analysis agree qualitatively with the experimental results.

A cartesian cut cell method for static body problems was presented in Part A (pp 47-56). Here, we extend the method to unsteady compressible flows involving arbitrarily moving bodies. The moving bodies are allowed to cross a stationary background cartesian mesh. So problems such as mesh distortion and/or restrictions on body motion which may affect other mesh approaches do not occur.

A MUSCL-Hancock finite volume scheme has been modified for moving boundary problems. The upwind fluxes on the interfaces of static cells are updated using an HLLC approximate Riemann solver and an exact Riemann solution for a moving piston is used to update moving solid boundaries (faces). A cell merging technique has been developed to maintain numerical stability in the presence of arbitrarily small cut cells and to retain strict conservation at moving boundaries.

The method has been validated against some well-known two dimensional test problems and applied to practical examples involving an exploding pressure vessel and a store release into a Mach 1.5 stream.

The Bell-Boeing V-22 Osprey tiltrotor is a unique aircraft capable of landing vertically, flying at speeds in excess of 300 KTAS (knots, true air speed), with the added feature of folding the rotor and wing for deployment from ships for US Navy, Marines, and SOF operations. During the development of the V-22, including 1200 hours of flight testing, a number of significant technology and design development challenges were encountered and overcome. This paper presents an overview of many of these challenges. It also reviews the integrated approach to testing to be used in the next phase of the flight test programme, and describes the changes that have been implemented to improve flight test productivity.