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To advance the design of heat exchanged gas turbine propulsion aeroengines utilising experience gained from early development testing, and based on technologies prevailing in the 1970‐2000 time frame.

With emphasis on recuperated helicopter turboshaft engines, particularly in the 1,000 hp (746 kW) class, detailed performance analyses, parametric trade‐off studies, and overall power plant layouts, based on state‐of‐the‐art turbomachinery component efficiencies and high‐temperature heat exchanger technologies, were undertaken for several engine configuration concepts.

Using optimised cycle parameters, and the selection of a light weight tubular heat exchanger concept, an attractive engine architecture was established in which the recuperator was fully integrated with the engine structure. This resulted in a reduced overall engine weight and lower specific fuel consumption, and represented a significant advancement in technology from the modified simple‐cycle engines tested in the late 1960s.

While heat exchanged engine technology advancements were projected, there were essentially two major factors that essentially negated the continued study and development of recuperated aeroengines, namely again as mentioned in Part I, the reduced fuel consumption was not regarded as an important economic factor in an era of low‐fuel cost, and more importantly in this time frame very significant simple‐cycle engine performance advancements were made with the use of significantly higher pressure ratios and increased turbine inlet temperatures. Simply stated, recuperated variants could not compete with such a rapidly moving target.

Establishing an engine design concept in which the recuperator was an integral part of the engine structure to minimise the overall power plant weight was regarded as a technical achievement. Such an approach, together with the emergence of lighter weight recuperators of assured structural integrity, would find acceptance around the year 2000 when there was renewed interest in the use of more efficient heat exchanged variants towards the future goal of establishing “greener” aeroengines, and this is discussed in Part III of this paper.

Turboshaft engines usually include one centrifugal compressor due to its high-pressure ratio, stability and compactness. Many designers rely on positive raking to decrease tip gap flow and therefore losses. However recent optimization studies revealed geometries contradicting this canonic view. Hence, this paper aims to investigate how the rake angle alone can influence performance and to which extent.

A turboshaft representative impeller was chosen and altered for null and +/−30° rake angles. Menter's shear stress transport model is used for steady computational fluid dynamics simulations, sweeping the nominal speedline at various tip clearances. Backsweep distribution is identical in all cases, isolating rake influence.

Pressure ratio was lowered for the both positively and negatively raked blades, but through distinct aerodynamic mechanisms. Although the flow through the tip gap was lower for the positive rake, this is due to lower blade loading. Splitter comparison reveal that these effects are more pronounced in the radial regions.

Some of the findings may extend beyond turboshaft engines, into turbochargers, home appliances or industrial blowers. However, all extrapolations must consider specific differences between these applications. Turboshaft compressors designers can benefit from this study when setting up their free parameters and penalty functions in the early concept stages.

Only few similar studies can be found in the literature to date, none similar to turboshaft applications. Also, this impeller is designed to eliminate leading edge shocks and suction side boundary layer separation, which makes it easier to isolate the tip gap flow effects. The authors also provide a framework on which semi-empirical design equations can be further developed to incorporate rake into 1D design tools.

This paper seeks to evaluate the potential of heat exchanged aeroengines for future Unmanned Aerial Vehicle (UAV), helicopter, and aircraft propulsion, with emphasis placed on reduced emissions, lower fuel burn, and less noise.

Aeroengine performance analyses were carried out covering a wide range of parameters for more complex thermodynamic cycles. This led to the identification of major component features and the establishing of preconceptual aeroengine layout concepts for various types of recuperated and ICR variants.

Novel aeroengine architectures were identified for heat exchanged turboshaft, turboprop, and turbofan variants covering a wide range of applications. While conceptual in nature, the results of the analyses and design studies generally concluded that heat exchanged engines represent a viable solution to meet demanding defence and commercial aeropropulsion needs in the 2015‐2020 timeframe, but they would require extensive development.

As highlighted in Parts I and II, early development work was focused on the use of recuperation, but this is only practical with compressor pressure ratios up to about 10. For today's aeroengines with pressure ratios up to about 50, improvement in SFC can only be realised by incorporating intercooling and recuperation. The new aeroengine concepts presented are clearly in an embryonic stage, but these should enable gas turbine and heat exchanger specialists to advance the technology by conducting more in‐depth analytical and design studies to establish higher efficiency and “greener” gas turbine aviation propulsion systems.

It is recognised that meeting future environmental and economic requirements will have a profound effect on aeroengine design and operation, and near‐term efforts will be focused on improving conventional simple‐cycle engines. This paper has addressed the longer‐term potential of heat exchanged aeroengines and has discussed novel design concepts. A deployment strategy, aimed at gaining confidence with emphasis placed on assuring engine reliability, has been suggested, with the initial development and flight worthiness test of a small recuperated turboprop engine for UAVs, followed by a larger recuperated turboshaft engine for a military helicopter, and then advancement to a larger and far more complex ICR turbofan engine.

Interest is currently being expressed in heat exchanged propulsion gas turbines for a variety of aeroengine applications, and in support of this, the aim of this paper is to evaluate the relevance of experience gained from development testing of several recuperated aeroengines in the USA in the late 1960s.

Technology status, including engine design features, performance, and specific weight of recuperated propulsion gas turbines based on radial and axial turbomachinery, that were development tested in the power range of about 300 to 4,000 hp (224 to 2,984 kW) is discussed in Part I.

A successful flight worthiness test was undertaken in the USA of a helicopter powered solely by a recuperated turboshaft engine and this demonstrated a specific fuel consumption reduction of over 25 percent compared with the simple‐cycle engine. However; in an era of low‐fuel cost, and uncertainty about the long‐term structural integrity of the high‐temperature heat exchanger, further development work was not undertaken.

The gas turbines tested were based on conventional simple‐cycle engines with essentially “bolted‐on” recuperators. Optimum approaches to minimize engine overall weight were needed in which the recuperator was integrated with the engine structure from the onset of design, and these are discussed in Part II.

Based on engine hardware testing, a formidable technology base was established, which although dated, could provide insight into the technical issues likely to be associated with the introduction of future heat exchanged aeroengines. These are projected to have the potential for reduced fuel burn, less emissions, and lower noise, and recuperated and intercooled turboshaft, turboprop, and turbofan variants are discussed in Part III.

THE RS.360 was conceived in the mid 60's as a turboshaft engine, for helicopter application and operation in the 1970's. It is one of three engines in the Anglo‐French Helicopter Agreement ratified in 1967. This agreement covered three helicopters — a medium, the SNIAS Puma with Turmo IIIC4 engines, a SNIAS LOH Gazelle with Astazou IIIN and a utility helicopter, the Westland Lynx. The RS.360 was chosen as the power plant for this helicopter and the design has been closely tailored to suit this aircraft. It is of modern design, but using ‘state of the art’ technology with the objective of providing a reliable, easily maintained power plant with particular emphasis on multi‐engine applications. A two spool gas generator layout was chosen for overall mechanical simplicity and to provide a good acceleration characteristic over the whole speed range. This arrangement eliminated the need for air blow‐off or sophisticated variable geometry. An important feature of the engine is the use of a modular concept permitting the replacement of major assemblies without the need for sophisticated equipment. Inherent in the design is the potential for adequate power growth. Provision has been made for extensive condition monitoring facilities in addition to normal flight requirements, ie intrascope access, magnetic chip detectors and SOAP.

The purpose of the present paper is to obtain the capability of designing modern rotorcrafts with enhanced accuracy and reliability.

Among the existing rotorcraft design programs, an appropriate program was selected as a baseline for improvement. It was based on a database comprising conventional fleets of rotorcrafts. The baseline program was not robust because it contained a simple iteration loop, which only monitored the gross weight of the aircraft. Therefore, it is not accurate enough to fulfill the quality and sophistication of a conceptual design framework useful for present and future generations of rotorcrafts. In this paper, the estimation formulas for the sizing and weight of the rotorcraft subsystem were updated by referring to modern aircraft data. In addition, trend curves for various turboshaft engines available these days were established. Instead of using the power estimation algorithm based on the momentum theory with empirical corrections, blade element rotor aerodynamics and trim analysis were developed and incorporated into the present framework. Moreover, the simple iteration loop for the aircraft gross weight was reinforced by adding a mathematical optimization algorithm, such as a genetic algorithm.

The improved optimization framework for rotorcraft conceptual design which has the capability of designing modern rotorcrafts with enhanced accuracy and reliability was constructed by using MATLAB optimization toolbox.

The optimization framework can be used by the rotorcraft industries at an early stage of the rotorcraft design.

It was verified that the improved optimization framework for the rotorcraft conceptual design has the capability of designing modern rotorcrafts with enhanced accuracy and reliability.

The Small Engine Division of Rolls‐Royce (1971) Limited, situated at Leavesden, Hertfordshire some twenty miles north of London is the smallest of the Company's three aero‐engine manufacturing plants and employs about 3500 people.

IN OCTOBER 1942, the first US jet aircraft — the Bell P—59 Airacomet — took to the air powered by twin General Electric 1—A jet engines. This flight marked the birth of the Jet Age in the United States. During the 1940's, General Electric continued to play the leading role in the development of flight propulsion technology with a series of high‐performance designs, based on the J47. In the 1950's, the J79 and J93 engines embraced variable stator technology for optimized performance. The air‐cooled GE1 engine family of the 1960's built upon this experience, permitting improved thrust‐to weight ratios, and marked the beginning of truly economical jet propulsion systems. Military, Commercial, Marine and Industrial powerplants in the 1970's — based on TF39 and CF6 high bypass turbofan technology — will develop more propulsion power at lower costs, with less maintenance than any other previous generation of engines. And these same GE engines incorporate no smoke and low noise designs to help make tomorrow's aircraft better neighbours.