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.

In the early 1970s considerable attention was being given to the possibilities of short-haul aircraft aimed at operating into or near city centres. There was a belief that suitable airfield sites could be found if field lengths no greater than about 2000 ft-3000 ft were required into which the aircraft could operate quietly with steep climb out and approach.

For the past 50 years civil air transport has been dominated by the combination of the gas turbine and the B-47 airframe configuration. Both these machines owed their creation to WW2 and both represented revolutionary concepts at the time. Oyer the years, the engine and the airframe have undergone considerable detailed refinement. The changes have resulted in very impressive improvements in performance. However, it is now very clear that both are close to some well defined asymptotic limits that bound their performance. The result is that further potential improvements are smaller in percentage terms, become harder to achieve technically and, consequently, come at increasingly higher cost.

A hugely successful global industry has developed around this standard configuration. Increasingly, more and more companies use aircraft and, as their businesses grow, the demand for aircraft and air transport increases. However, the ability of the environment to absorb this growth is beginning to be questioned and new ideas are required if aviation is to be regarded as an unbounded and a truly sustainable activity.

Installation effects on the S-duct intakes of a high power turboprop were investigated by comparing three different nacelle configurations as the channel wing, over-the-wing and conventional tractor wing using computational fluid dynamic methods. The interaction of the propeller, the wing, and the nacelle on the scoop type turboprop intakes were identified in terms of recovery and distortion on the aerodynamic interface planes. An actuator disc model was used to simulate the propeller downstream effects. The channel wing installation of S-ducts showed approximately 2% higher total pressure recovery than the other configurations. The channel wing configuration with a wrap-around S-duct experienced a 2% lower circumferential distortion intensity, whereas over-the-wing had an increase by 3% in comparison to the tractor wing. It was also observed that the S-duct type had a greater influence on the radial distortion intensity than the installation type. The swirl coefficient distribution showed that multiple peaks occurred as opposed to swirl results based on a single sector of the aerodynamic interface plane. The difference between swirl peaks of shaft penetration S-duct was reduced by 14% when installed on channel wing, whereas the peaks of wrap-around experienced 90° phase shift, and their difference showed 10% increase on channel wing.

Along many flight corridors, bodies of water serve as preferred emergency landing options. Thus, relevant scenarios must be investigated to improve aircraft crashworthiness in the event of an impact landing on water. Enhancing the damage tolerance of aircraft structures through repetitive experiments can, however, prove highly uneconomical. Such large-scale trials can be influenced by many factors of uncertainty adversely affecting the quality of the results. Therefore, the work presented in this study focuses in particular on evaluating a computational methodology perfected for aircraft water ditching using Coupled Lagrangian-Eulerian (CLE) that allows detailed prediction of structural response of a verified deformable fuselage section during such events. Validation of the fluid-structure interactive (FSI) strategy developed is conducted, thoroughly comparing the method against the analytical and experimental results of multiple wedge drop tests. Finally, the validated FSI strategy is applied to a high-fidelity fuselage section model impacting water to simulate and assess a realistic ditching scenario.

This paper presents the results of an interferometric study of the shock/boundary-layer interaction at a compression corner. The boundary layer upstream of the corner is a non-equilibrium state having been disturbed by injection through a porous surface. As well as being relevant to the design of transpiration-cooled surfaces, the results show many similarities to the interaction in an adverse pressure gradient. The changes in the nature of the interaction as the approaching profile is distorted is fully explored and the results should provide a useful test case for computational fluid mechanics. The results are also used to evaluate the accuracy of interferometric techniques.

Typical performance characteristics are presented for a cascade of two-dimensional bluff aerofoils employing circulation control from a tangential blowing jet. The cascade was tested in various configurations at a freestream Mach number of 0·3 and Reynolds number of approximately 1·5 x 105. Lift and drag coefficients and stream deflection are related to the blowing jet momentum coefficient. The expressions employed for lift and drag eliminate the direct effect of the blowing jet on measured performance, which was found to be influenced strongly by vortex shedding at low jet blowing rates.

For the configurations employed, tests showed that the economical application of circulation control on an uncambered blade could produce flow turning angles up to 20°.

The object of this paper is to give a brief account of the development of the study of hypersonic flow over the past three decades, and to stress some of the ways in which hypersonic flow differs from supersonic flow. In some of the early work on the subject a wrong emphasis was given to the study of flows with attached shock waves and to the nature of the shock boundary layer interaction problem, and it is only quite recently, mainly as a result of experimental work, that it has become possible to obtain a balanced picture of the important features of hypersonic flow fields. In the present paper real gas effects are not taken into account; in general, the inclusion of these effects will not markedly alter the picture given.

This paper is concerned primarily with some of the practical difficulties encountered in connection with the prediction of kinetic heating temperatures. Attention will be concentrated upon methods for estimating temperatures and heat transfer rates for practical aircraft designed to fly at Mach numbers up to about five.

One factor common to all kinetic heating calculations is the variation of temperature through the thickness of the boundary layer, with consequent variation of viscosity. At Mach numbers above about 3, these temperature variations also lead to considerable variations of other properties of air—which are commonly assumed to remain constant—even in classical compressible flow aerodynamics. These factors complicate the aerodynamic equations.

The estimation of the relative orders of the coefficients in the linear differential equations of longitudinal small perturbation of aircraft is used to analyse the classical simplified theory for the short- and long-period modes by the singular perturbation methods. The leading terms of the asymptotic expansions give the eigenvalues for short- and long-period modes with fairly good accuracy, except that the damping part of the eigenvalue for the long-period mode has relative error of the order of 10%. These approximate eigenvalues are improved by the method of strained co-ordinates which accounts for most of the differences. The dynamic response due to initial disturbances of velocity, angle of attack, and angle of pitch are also derived.

In space, structures encounter various severe environments, including severe thermal conditions. The signal level of the radio wave from the Large Deployable Reflector (LDR) mounted on the Engineering Test Satellite-VIII (ETS-VIII) was observed to change during an Earth eclipse. This phenomenon was assumed to be caused by the thermal deformation of the LDR. Therefore, in this study, a means to suppress the thermal deformation is proposed and demonstrated by focusing on the internal force generated at the springs used to deploy the antenna. According to the numerical results obtained from finite element analyses, the thermal deformations at all apices that support the reflectors were suppressed at a high correction rate by adjusting the coefficients of thermal expansion in the structural members and by controlling spring forces differently in four areas depending on the distances from the constraint point.

A Lean Premixed Prevaporised (LPP) low-emission combustor with a staged lean combustion technology was developed. In order to study cold-flow dynamics in the LPP combustor, both experimental tests using the particle image velocimetry (PIV) to quantify the flow dynamics and numerical simulation using the commercial software (FLUENT) were conducted, respectively. Numerical results were in good agreement with the experimental data. It is shown from the observation of the results that: there is a Primary Recirculation Zone (PRZ), a Corner Recirculation Zone (CRZ) and a Lip Recirculation Zone (LRZ) in the LPP combustor, and the exchanges of mass, momentum and energy between pilot swirling flow and primary swirling flow are contributed by the velocity gradients, and the shear flow is transformed into a mixing layer exhibiting the higher Reynolds stresses, which suggests the mixing process is strictly affected by the Reynolds stresses.

This paper describes the first work undertaken to see the flows of oil, kerosene, water or any other hydrogenous fluid within a steel casing. In the specific case described, the performance and movement of the oil in the scavenge system of the Rolls-Royce Gem helicopter gas turbine was clearly seen with the engine operating at a range of test conditions.

The work was conducted using cold neutrons from the DIDO nuclear reactor at AERE, Harwell, England, and this was the first time that such work has been attempted. The experiment was a complete success and provided answers to developmental problems in the oil system.

In the paper, the preliminary work on the ‘Herald’ reactor at AWRE, Aldermaston, leading to this breakthrough in technology is also described.

While the use of blade lacing wires or shrouding is not altogether common in gas turbine practice, occasions may arise when their use is desirable, and the purpose of this note is t o present a method whereby the vibrational characteristics of a number of blades wired together may be conveniently found.

Accurate prediction of the response of low-aspect ratio, flexible aircraft requires correspondingly accurate modeling of the aircraft itself and of the aerodynamic forces, both respectable problems. Assuming that the wing can be modeled as a nonuniform plate, the discretisation process of choice is the finite element method (FEM), which demands a very large number of degrees of freedom for good accuracy. Moreover, accurate modeling of the aerodynamic forces acting on the aircraft suggests the use of computational fluid dynamics (CFD), which requires the use of an extremely large number of variables. On the other hand, feedback control design for the aircraft demands an aircraft model of relatively small order, so that the dimension of the FEM and CFD models must be reduced drastically. Based on physical considerations, reasonably accurate model reductions can be achieved, but a problem remains because the FEM and CFD grids are likely to differ from one another. It is shown in this paper how to achieve desirable model reductions for both the FEM and CFD and how to integrate the aerodynamic forces into the aircraft state equations. A numerical example demonstrates how the theory can be applied to the flight of a flexible aircraft. The analytical/computational approach developed here should permit parametric studies ultimately resulting in a reduction in the time required for aircraft design and flight testing.

For the safety problems caused by the limited landing space of the deck during the arresting process of the carrier-based aircraft, a dynamic model of the carrier-based aircraft’s landing and arresting is built. Based on the batch simulation method, the lateral dynamics safety envelope of the aircraft during the arresting was defined, and the dynamic response of the key points in the envelope during the arresting process was investigated. Subsequently, the influence of engine thrust and aircraft quality on the arresting safety envelope was studied based on reasonable safety evaluation indicators, and the safety status envelope of the deck arresting was given. Then, the particular Hamilton-Jacobi partial differential equation is used to obtain the lateral dynamics safety envelope of the carrier-based aircraft in the process of landing and arresting by backward inversion. Results indicate that engine thrust and landing quality have little effect on the yaw angle in the arresting safety boundary during the arresting. Additionally, with the engine thrust and landing quality increase, the maximum safe off-centre distance gradually decreases, and the safety boundary decreases accordingly. During the phase of landing glide, the engine thrust and quality have little effect on the maximum safe eccentric distance. When the engine thrust is increased by 40%, the maximum safe yaw angle is reduced from 0.3°, and the safety boundary is reduced by 4.2%. When the aircraftquality increases by 40%, the maximum safe yaw angle is reduced by 0.4°, and the safety boundary is reduced by 2.8%. The findings of this paper can provide framework for the research on theaircraft-to-carrier dynamic matching characteristics of the carrier-based system, and is of great significance to the research on improving the safety of the carrier-based aircraft landing arresting.

The integral equations for the pressure distribution on an oscillating thin wing having a prescribed velocity distribution in a supersonic flow are developed.

An alternative method is described for solving the problem of a three-dimensional jet-flapped wing with a high aspect-ratio. This method is similar to the lifting-line theory of Kerney6 or Tokuda7, but differs in that the method of matched asymptotic expansions is applied to an integral equation, derived from the lifting surface theory, rather than a partial differential equation. The advantage of the present method over those used previously is that the necessary outer solutions are obtained directly; it is not necessary to rely upon physical insight or considerable ingenuity. The final results are different from those obtained by the previous authors; it is shown that the present result is correct, by noting some errors in the earlier theories.