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The purpose of this paper is to analyze the effects of the PRD geometric parameters, including the area and aspect ratio, on the discharge and force characteristics of pressure relief process under various plenum compartment pressures and Mach numbers.

Under various plenum compartment pressures and Mach numbers, the effect of the area and aspect ratio on the discharge and force characteristics of the PRD are numerically investigated via a three-dimensional steady Reynolds-averaged Navier–Stokes equations solver based on structured grid technology.

When the aspect ratio remains constant, the discharge coefficient C_D, thrust coefficient C_T and moment coefficient C_M are not affected by the PRD. When the area is constant, the aspect ratio dramatically impacts the discharge and force characteristics because the aspect ratio increases, the discharge coefficient C_D of the PRD decreases, and the thrust coefficient C_T and the moment coefficient C_M both increase. When the aspect ratio is 2, the discharge coefficient C_D decreases by 14.7 per cent, the thrust coefficient C_T increases by 10-15 per cent, and the moment coefficient C_M increases by 10-23 per cent compared with when the aspect ratio is 1.

This study provides detailed data and conclusions for nacelle PRD researchers and actual engineering applications.

On the basis of considering the influence of operating conditions on the discharge and force characteristics of the nacelle PRD, the impact of geometric parameters, including the area and aspect ratio on the discharge and force characteristics is comprehensively considered.

THE Trident IE fuel system, designed to operate on cither kerosene or JP.4, has a straightforward layout with few controls. Five integral tanks (FIG. 1), comprising four in the wings and one in the centre section, give a total of 5,880 Imp. gall, of which 2,000 Imp. gall, are contained in the centre tank. (Total fuel capacity of the Trident 1C is 4,960 Imp. gall, with 1,160 Imp. gall, in the centre tank.) Each wing inner tank has slightly more than twice the capacity of the outer.

Aerodynamic characteristics of engine side air intakes for a lightweight helicopter are investigated aiming to achieve an efficient engine airframe integration.

On a novel full-scale model of a helicopter fuselage section, a comprehensive experimental data set is obtained by wind tunnel testing. Different plenum chamber types along with static side intake and semi-dynamic side intake configurations are considered. Engine mass flow rates corresponding to the power requirements of realistic helicopter operating conditions are reproduced. For a variety of freestream velocities and mass flow rates, five-hole pressure probe data in the aerodynamic interface plane and local surface pressure distributions are compared for the geometries.

In low-speed conditions, unshielded, sideways facing air intakes yield lowest distortion levels and total pressure losses. In fast forward flight condition, a forward-facing intake shape is most beneficial. Additionally, the influence of an intake grid and plenum chamber splitter is evaluated.

The intake testing approach and the trends found can be applied to other novel helicopter intakes in early development stages to improve engine airframe integration and decrease development times.

Details of the flying controls, electric and electronic, ice protection, environmental and fuel systems. The design principles of the flying control system for the Series 2 Skyvan have been retained in the Series 3 aircraft. The original plan to have interconnection between the aileron and flap control circuits giving aileron droop of approximately half the flap deflection was shelved on production versions of the Series 2 aircraft due to the performance of the slotted flaps being better than estimated.

In this article the requirement leading to the design is discussed and this is followed by a general description of the aircraft and its operating efficiency. The accommodation is described, with particular reference to the flight deck and equipment. The various systems and installations are next reviewed; followed by a detailed description of the various structural components, which, in many cases, have been designed round these systems. The development work that has been done in order to allow design and construction to be completed will be described in some detail in a later article.

ON August 10, 1949, the Avro C102 jet transport, now better known as the Jet‐liner, made its first flight.

THE AIRCRAFT MANUFACTURER can provide helpful answers to the two most important questions of the communities affected by air transportation. Aircraft will be built which do not intrude on the every day existence of communities. Aircraft will be built whose airports only require tens of acres of land instead of thousands of acres. The DHC‐7 will be the first of them.

The propulsion system integration of a turboprop aircraft, which has been developed for the basic trainer, was performed. The proper turboprop engine was selected among worldwide existing engines by the specific developed engine selection technique and trade‐off studies such as customer’s request for operational capability (ROC), propulsion system parameters, performance analysis with engine installed effects, future growth potential, integrated logistic support (ILS), maintainability, interfaces with the airframe, etc. The chin type air inlet with the plenum chamber was designed in consideration of the inclined configuration to minimize the propeller swirl effect, the inertial separation bypass device to reduce FOD, and the super‐ellipse and NACA‐1 profile lip to maximize the ram recovery. The air inlet was analyzed by a higher‐order source panel method considering propeller wake. The exhaust duct was designed through internal cross‐section area determination to maximize the cruising power as well as external configuration to maximize the effective power, to minimize the aerodynamic drag and to minimize the cockpit contamination by the exhaust gas. The proper oil cooler for the selected turboprop engine was determined with cooling requirements and the oil cooling inlet duct with NACA configuration was designed. The test of the propulsion system including the installation performance test with the effects of the air inlet, the exhaust duct, the propeller and the nose fuselage configuration was performed in the test cell.

A design study is carried out for an air cushion vehicle based on a specification giving it a performance comparable with cars of similar power and size, and compatible with current road conditions. Preliminary examination of requirements leads to the adoption of a plenum chamber lift system and ducted fan propulsion, powered by a single piston engine. Care is taken to provide the vehicle with the responsive and accurate control necessary on the road. Performance calculations are carried out, enabling it to be compared with wheeled road vehicles. It is found that its cruising speed and fuel economy are better than its wheeled counterpart, due to the low resistance to motion offered by the air cushion, while its turning performance is definitely inferior.

Owing to the inadequacy of conventional laboratory windshield qualification testing methods, it is not unusual for unforeseen problems to develop in the real world of flight service. Aircraft windshields are normally subjected to static, material, environmental, and, in some instances, to limited service condition testing before being incorporated into prototype hardware. In‐service monitoring is used to compare the prototype to previous designs. Determining the relative effectiveness of corrective “fixes” or design improvements by in‐service testing can take months or even years of calendar time. To overcome these prob‐lems, Sierracin has developed and constructed an advanced Windshield Flight Environment Simulator (WFES), that duplicates the total operating environment of electrically‐heated, pressurized windshields, on a highly compressed time scale.