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An important component within the jet engine in terms of vibration and high cycle fatigue (HCF) is the blade. This is the component where continuously higher demands on weight and loading are being made. As a consequence of this, there has been a growing interest in developing both numerical methods and instrument technology for blade HCF measurement. This growing interest has also been attributed to changing attitude within the military and aerospace industry, which has tended towards driving down costs and lengthening the engine's life span. Many development technologies have been reported. One of which, is the development of a non-intrusive system for measuring blade vibratory stress. Research in non-intrusive techniques for the measurement of blade vibration has been ongoing since the early 1970' s. The aim of which, has been to replace the conventional method, using strain gauges and slip rings, with an improved system based upon non-intrusive type instrumentation such as optical or capacitance probes. One such approach is known as tip-timing. Tip-timing is a technique used to measure blade vibration using non-contact probes located around the engine casing. Many tip-timing techniques have been developed over the years, but there still remain significant problems associated with the approach. Such problems include sensitivity to noise and the high number of probes required. The development of two tip-timing methods known as the Autoregressive (AR) method and the Two Parameter Plot (2PP) method has recently been published in the open literature. This thesis describes the work done to experimentally test these two techniques. During the course of this work, an experimental optical tip-timing test facility was built. This included purpose-built optical tip-timing instrumentation, a tip-timing data acquisition system, and a post processing system incorporated into the Cranfield University low speed compressor facility. Experimental testing of the Autoregressive method and the Two Parameter Plot method was carried out using a controlled test environment, representative of a real engine. An analysis of the two methods was conducted using data from a comprehensive range of frequencies and RPM speeds. The results were then compared with previously published numerical results and the two algorithms were evaluated in terms of replacing the conventional strain gauge method. Testing of the AR method presented some interesting findings, with acceptable results produced at low rotational RPM speeds. However, as the rotational speed was increased, the accuracy of the results deteriorated. This type of result had not be highlighted in previous work. The 2PP method performed relatively well when using data sampled from the smaller 16 Engine Order (EO) response. However, this was not repeated when using the larger 72EO data. Additionally, this type of result had not been shown in previously published work. Overall, it was concluded that the issues associated with the frequency measurements should be remedied and a technique for measuring Multiple-Degree-of-Freedom responses should be explored before tip-timing techniques can be considered as a replacement to the strain gauge approach.

No description available.

No description available.

No description available.

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To meet increasing and often conflicting demands on performance, stability, fuel consumption and functionality, modem jet engines are becoming increasingly complex. Improved compressor performance is a major factor in this development process. Optimum compressor efficiency is achieved in operating regions close to flow instability. Surveying basic concepts and control methods of compressor instabilities, an overview of the fundamentals of surge and rotating stall is presented. To maximise the potential of an aero gas turbine compression system, it is proposed to use more advanced control strategies, such as multi variable control. Multivariable control may offer the prospect of lower safety margin requirements leading to greater compressor efficiency. Alternatively, it may result in more agility in combat through improved engine responses and prolonged engine life. A multivariable control technique is proposed and tested on a Rolls-Royce three-spool high bypass ratio turbofan engine. Since elements of the 2x2 system can be represented by linear third order models, a muItivariable PID controller will be sufficient provided the design requirements are not too rigorous. To have a simple and efficient design, a systematic decentralised PI (PID) control design strategy is developed. Decoupling a given 2x2 process by a stable decoupler, the elements of the resulting diagonal matrix are approximated by first (second) order plus dead time processes using the proposed model reduction techniques. Then, SISO controllers are designed for each element using the developed tuning formulae. Any practical design method should be simple, easy to apply, flexible, generic or extendable, and applicable to complex control schemes to fulfil more demanding control requirements. It will be advantageous if the design algorithm can also directly address the design requirements, be repeatable for any control objective, constraint and category of processes, have a design parameter, and can consider any number of objectives and constraints. Formulating the PI (PID) control design problem as an optimisation problem, a non-dimensional tuning (NDT) method satisfying the above-mentioned design properties is presented. For a given first (second) order plus dead time process, the NOT method is used in conjunction with either a single-objective or a multi-objective optimisation approach to design PI (PID) controllers satisfying conflicting design requirements. In addition, considering load disturbance rejection as the primary design objective, a simple analytical PI tuning method is presented. The design problem is constrained with a specified gain or phase margin. Compared to the corresponding conventional SISO controller, it is demonstrated that the resulting decentralised controller considerably improves the overall surge risk to the engine during the transient manoeuvres while maintaining similar thrust levels. Due to non-linearity of jet engine models, gain scheduling is necessary. Designing decentralised controllers at various operating points, the gain-scheduled controller accommodates the non-linearity in engine dynamics over the full thrust range.

A computer model of the J-85 gas turbine engine has been used in a investigation for potential benefits in performance simulation, arising from the adoption of compressor stage stacking models. The eight-stage, axial flow compressor of the engine was simulated by means of two separate stage stacking methods: the Howell program and the projects code, the latter being developed on the basis of the pre-existing J-85 complete engine simulation model. Results obtained from the use of the Howell program generally support its suitability for the prediction of single stage and overall compressor performance. This includes the capability of the code to render the effect from the Variation of specific design parameters and from the incorporation of variable geometry. Certain weaknesses originating from the empirical nature and the logic of the program are identified. Recommendations for the assessment of the validity of produced results, a well as modifications for the improvement of the code, are proposed. The projects stage stacking code was tested with sets of stage characteristics derived from different techniques. The produced results depend on the topology and shape of the utilised characteristics. A special feature embedded in the program's logic denotes the tendency of individual stages to work within their stall areas or below the hypothetical choke points of the corresponding characteristics. Although in certain instances the predicted overall compressor performance appears satisfactory, analysis of results in individual stages indicates the need for improvement of the code, in order to obtain a closer approach to the physical mechanism and the associated limitations of stage matching. This is also true for the prediction of variable geometry effects on the overall compressor performance. The complete engine simulation model incorporating the projects stage stacking code provides generally satisfactory results, a compared to those produced by a equivalent model of the Cranfield's Turbomatch Scheme. Specific problems encountered are attributed to certain weaknesses of the embedded stage stacking code, which is susceptible of improvement as already reported.

To remain competitive, gas turbine manufacturers must aim for continuingly improved engine efficiencies and thrust-to-weight ratio. This has resulted in the design of gas turbines with increased turbine entry temperature (TET). Thermal barrier coatings (TBCs) are the most promising systems, which thermally protect engine components and allow their use at higher engine gas temperature by potentially reducing metal surface temperature by up to 150°C. The TBC system consists of a metallic bondcoat and a thermally insulating strain-tolerant ceramic top coat. The bondcoat is a critical part of the system; its failure has a major impact on the lifetime of the TBC. The purpose of this work is the development of a novel and innovative bondcoat with reduced weight, also called "low-mass" bondcoat. This new class of bond coat consisted of a thin (2.5 to 8 J..lm thick) coating containing successive layers (from 9 to 163) of aluminium and platinum. The layers react with one another exothermically by diffusion after a subsequent heat-treatment at a relatively low temperature (700°C), to form an intermetallic bond coat. In this thesis, the manufacture and optimisation of the low-mass bond coat TBC are presented and discussed. Deposition prerequisites along with good deposition practice were defined in order to produce successfully the low-mass bond coat in a clean environment. Stable working parameters were established, among which a roughness working window, as the substrate initial roughness appears to be a key parameter for coating adherence. The structure of the individual as deposited layers were characterised, which allowed to determine the surface temperature during deposition (between 150°C and 350°C). This was well below the temperature above which the exothermic reaction is triggered (400°C). High-multilayered bondcoats (PtAI, PtAh, Pt2Ah stoichiometries) were successfully manufactured, characterised and integrated in a TBC system, among which the thinnest bond coat for THC ever made (51 layers for a 2.5 J..lm thick PtAh). The low-mass bond coat TBC system presented a singular structure consisting of a dense intermetallic layer overlaid by a composite structure of Ah03 precipitates within a (Ni,Pt)xAly matrix. Furthermore the TGO, thermally grown oxide, formed and grew with a typical equiaxed granular structure. This novel TBC system was tested along with commercial coatings under thermal cyclic oxidation, aiming to simulate the thermal cycles induced by the operating aircraft gas turbine .. Regarding to the thickness and the aluminium reservoir of the low mass bond coats, the performances are outstanding, confirming the potentiality of this new type of TBC systems. A degradation mechanism was proposed based on FIB and SEM observations along with chemical analysis. The outstanding performance of the low mass bond coat TBC system is thought to be due to the very specific manufacturing process and its influence on the alumina scale growth under the TBC.