Performance data are presented for eight annular diffuser families, a family being defined by the inner and outer cone angles and the inlet hub-tip ratio. These results include profiles of velocity and total pressure across the inlet and outlet lanes as well as overall pressure recovery coefficients. Diffuser outer cone angles range from 10° to 60° and inner cone angles from 59° to -19°. Swirl angles were varied from 0° up to 44°. Each diffuser family was tested over a range of lengths and swirl angles resulting in over ninety different tests. A procedure for extending the non-swirling definition of effectiveness to swirling flows is presented and used to evaluate a limiting swirl angle for separation for free-vortex-type inlet swirl. Comparisons are made to several measured values of flow separation. Some experimental diffusers exhibited a violently unsteady hub stall which apparently consisted of rotating reverse flow regions.

The work reported here was an experimental study of the combined influence of a sandpaper rough surface and a stream turbulence upon laminar and transitional boundary layers. Data correlations were obtained using the previously determined parameter of an equivalent Nikuradse size of roughness. These correlations are derived for the development of the laminar layer in thickness and profile shape factor, for the onset and the length of transition and for the distribution of intermittency through transition. Further fundamental differences in flow behaviour from those in a rough pipe have been found to add to those already reported.

This paper reports on the experimental validation of a new model of slider-reefed ram-air parachute opening, using jump data collected by this investigator and others. The model is based on the parachute load equation of motion coupled with a new differential equation which describes the rate of change of the parachute's drag area. The model clearly identifies the important factors which determine the inflation peak load and opening time. It also features many scaling relationships useful for design work. Several predictions are shown to be design-independent. The experimental data used in this model validation were gathered during sport parachuting jumps performed by this author. Also used are data generated by the US Navy during tests of the MC-5, MT-1X/SL and MT-1XX parachutes. The experimental results and model predictions are found to agree within 20% accuracy over a wide range of opening speeds and deployment altitudes.

The unsteady flowfield in a twin side-by-side intake arrangement arising as a result of a surge in one of the engines has been studied numerically by solving the Euler equations. Unsteadiness was introduced by prescribing a pressure disturbance at the exit plane of one of the intakes. The two cases considered correspond to static ground-running of the engines and a flight Mach number of 0.6 with a single rapidly increasing exit plane pressure disturbance representative of a pop surge. The amplitude of the imposed pressure disturbance varied between 100% and 200% of the mean exit static pressure. In the cases considered, the results indicate that static pressure attenuation of the propagating hammer shock occurs upstream of the intake entry plane, resulting in a relatively weak rarefaction wave travelling down the adjacent intake. No evidence of increased dynamic flow distortion likely to lead to a complementary surge in the adjacent engine was observed. These conclusions have been confirmed by available test data thus demonstrating the value of CFD techniques for modelling complex unsteady transient flow phenomena of this type.

Potential function methods have been applied extensively to many terrestrial and space robot control problems. This paper extends these methods to multiple robot problems for on-orbit assembly. By considering a population of free-flying robotic vehicles in orbit, capable of both rotation and translation, the potential function method reduces complex assembly problems to sets of simple translation and rotational commands. This is achieved using a potential function incorporating both collision avoidance between the free-flyers and connection constraints in the assembled structure. The method also allows a subsumptive type control architecture with the flexibility to carry out many different assembly tasks simultaneously.

Previous experimental studies have shown that the development of flow behind the trailing edge of a lobed forced mixer is greatly influenced by the formation of large-scale streamwise vortices. The main objective of the present study is to establish quantitatively the presence and role of the streamwise vortices in the mixing layer in the near field of a lobed forced mixer trailing edge. The near field was defined as the region within five wavelengths (of the trailing edge profile) from the trailing edge. A two-stream mixing layer with a velocity ratio of 0.6 was generated with initially turbulent boundary layers and nominally two-dimensional flow. Mean flow and turbulence measurements were made on fine cross-plane grids across the wake region and at several streamwise locations using a two-component laser-doppler anemometer. Production of turbulent kinetic energy existed in most parts of the near field enhancing the overall mixing process. The streamwise vortex development underwent a three-step process by which it was formed, intensified and quickly dissipated towards the end of the near field. The distance between two rows of streamwise vortices of alternate signs within a lobe reduced with downstream distance causing the amalgamation and annihilation of the vortices. Analysing this movement of vortices, based on inviscid vortex dynamics, suggested that the normal vortices shed at the trailing edge had actually provided a squeezing (pinched off) effect to two adjacent rows of streamwise vortices within a lobe, forcing them to interact with each other.

The implications of a rise in the pressure drag of a wing near its junction with a fuselage are examined. Based on some data measured for a swept-wing-plate junction, in which the wing was of NACA 0015-profile normal to its leading edge and swept back at 20°, it is found that there is a potential penalty of order 1% of the total aircraft drag in the cruise condition.

The NH90 project formally started in September 1992 when France, Germany, Italy and The Netherlands agreed a contract with NH Industries on the design, development and qualification of an 8-9 tonne class helicopter which shall replace the present fleet of light transport and naval helicopters in the various military services of these states.

The NH90 comprises two versions that are based on a weapon system development specification for a Tactical Transport Helicopter (TTH) and a Nato Frigate Helicopter (NFH). The design of the NH90 follows the principle of having as many common basic systems and subsystems as possible in both versions, and to utilise a high level of avionics system integration to provide the capabilities to carry out the required tactical missions with a minimum of crew and crew workload.

Consequently, structure, dynamics, general vehicle, and flight control systems, core avionics, and mission equipment are principally common features for both weapon systems. The integration of two specific mission systems, on the other hand, provides the capabilities to employ both versions with a high rate of efficiency in their individual roles. The TTH tactical transport for air forces and armies, and the NFH as a ship embarked anti-submarine/anti-surface unit warfare (ASW/ASUW) for navies, both with the potential for a variety of other missions. The various subsystems of both mission systems, use modern equipment for acquisition of mission relevant information and data, their processing, transformation and display and provide automatic actions or advise to initiate adequate crew reactions.

Convective heat transfer data is frequently obtained from transient surface temperature measurements. Thin film resistance gauges, thermocouples, and thermochromic liquid crystals, are used in various situations to measure the surface temperature history. By assuming that uniform semi-infinite flat plate conditions apply, it is possible to express the instantaneous surface heat flux as an analytical function of the transient surface temperature(1). Various approaches can be used to account for the presence of multi-layered substrates and finite thickness substrate effects (Schultz and Jones(1); Doorly and Oldfield(2); Guo et al(3)), however, the effects of surface curvature are usually neglected.

Two-dimensional Euler equations are solved on unstructured triangular meshes using commonly available minicomputers. The driving algorithm is an upwind biased implicit cell-centred finite volume scheme. The spatial discretisation involves a naturally dissipative flux-split approach that sharply captures Shockwaves. The grid generation method uses a type of advancing front technique which triangulates a given set of points. To demonstrate the application, the steady flow results are presented for single and two component aerofoils. The unsteady results are obtained using a dynamic mesh algorithm for an aerofoil pitching harmonically about the quarter chord. The paper presents a description of the grid generation and movement algorithm and of the Euler solver, along with results and comparison that assess their capabilities.

Measurements have been made in the exit flows of the fan, diffuser, settling chamber and contraction of a low speed blower windtunnel driven by a centrifugal fan. A velocity deficit and streamwise vortex were observed at the diffuser exit when the fan was operating at less than about 0·9 of maximum efficiency but not when operating above. Unacceptably large unsteadiness (-0·7% rms) occurred at low speeds when the fan speed was increased from rest, but not when the fan speed was decreased from full speed. This unsteadiness was at the rotor frequency, and appears to have been caused by laminar boundary layer separation on one or more adjacent aerofoil blades, coupled with slight asymmetry in the rotor. Unsteadiness at the rotor frequency and of comparable magnitude has also been seen in another, larger fan, but was not eliminated in the same way. Attaining high quality flow is likely to require careful attention to the characteristics of individual fans.

The selective bleed turbofan is a two shaft, three compressor, variable cycle aircraft engine. Its bypass ratio (BPR) can be modulated to suit the operating flight conditions. At subsonic flight speeds it operates as a medium bypass turbofan (low pressure mode or LPM). It becomes a low bypass turbofan (high pressure mode or HPM) when flying faster and is capable of supersonic cruise without reheat (RH).

The aim of this paper is to describe the selection of the main engine design parameters in the HP mode and to outline the challenges to be met to enable the powerplant to operate in the other mode. The components have to perform satisfactorily over a wide range of varying conditions. The resulting demands are reflected in many ways, for example in the operating lines of the low pressure compressors and in the overall efficiency changes with increased bypass ratio. In the case of the present investigation the two bypass ratios selected are 0·3 for the high pressure mode and 0·7 for the low pressure mode. The latter was selected after an examination of a choice of bypass ratios ranging from 0·4 to 1·0. The change in variable stators required to achieve the LP mode cycle is indicated.

The effect of changing bypass ratio on engine performance and handling is shown. The main conclusion is that the design of this engine seems to be feasible within current technological capabilities and further investigation is encouraged because it appears to yield significant benefits.

Recent experience in the prediction and assessment of windtunnel wall interference is reviewed. Three areas of development are identified. The first deals with the influence of wall boundary layers on wall-induced velocities in solid-wall windtunnels; the second is concerned with preparations for determining wall interference in the European Transonic Windtunnel and the third relates to recent studies aimed at providing the basis for correcting data to free-air conditions.

What is of central importance in progressing new materials through from development to use is not technology but the creative efficiency of the team. What results is technology, both a competence and a confidence in handling the new material that makes the clarity of the demonstration of its potential a delight. It requires a holistic approach. Clear concepts require functionally direct embodiment, and equally expressive robust detailing. No amount of additional detail will retrieve a bad concept. Growing a team that has enough insight to tell the difference and the courage not to settle for less than the best is the challenge.

A methodology for the analysis of the performance of different aircraft configurations in the conceptual design phase is addressed. A scalar objective function, takeoff weight, is minimised by means of a numerical optimisation technique which takes into account the high number of geometrical parameters and the flight mechanics requirements involved in the problem. The study is confined to the cruise segment of the mission, and results are shown both for tail-aft and canard configurations. The above technique produces “equally” optimised configurations, thus giving a meaningful means to compare tail-aft and canard configurations.

A critical revision is made of the classic power-law expressions that describe the effect of the Prandtl number on laminar heat transfer and temperature recovery for Pr near one, by applying a perturbation analysis to the energy equation for constant-property flow. In this way, the Pr dependence is obtained as a systematic asymptotic expansion up to any desired order, which provides a more accurate and convenient procedure than classic curve-fitting methods. The effect of pressure gradient and flow sweep are examined for self-similar yawed wedge flows.

An expression is derived to predict the effect of taper ratio on the vortex drag of high aspect ratio, unswept, planar wings. The expression was formulated by assuming that imposing an elliptic spanwise load distribution on the tapered planform, would result in a leading edge thrust excess, which relates to the increase in vortex drag above the theoretical minimum. Comparisons of this expression with results from lifting line theory show encouraging agreement.

The transonic facility of DLR Göttingen (TWG) has been modernised with respect to an improved flow quality and flow simulation, as well as to the logistics of exchangeable test sections, operational reliability and productivity. It is presently equipped with a Laval nozzle for supersonic flow measurements, a perforated test section (6% open, 60° slanted holes) for transonic measurements, and a two-dimensional adaptive test section for subsonic measurements. The latter can perform two-dimensional adaptive tests on wing profiles, using the Cauchy integral formula for wall adaptation, as well as two-dimensional wall adaptation for three-dimensional models, utilising the Wedemeyer/Lamarche procedure, also known as the VKI method. For both cases, it is shown that the wall adaptation was successful. The 3D force results compare quite well to test results from the perforated test section, as well as to nominally interference-free results. The pressure distribution from the wing profile tests compare quite well to theoretical calculations. In the course of windtunnel modernisation, the transonic facility of DLR Göttingen has been equipped with the new software system DeAs for data acquisition and control of experimental facilities. In this environment the wall adaptation programs had to be implemented. Before adjusting the computed wall shape, it has to be controlled in the loop for sufficient accuracy and tolerable bending stresses. A simplified finite element method — also taking into account the pressure loads at the wall — is used for this purpose. The simplified approach has been checked against more detailed computations using ANSYS and NASTRAN.