Volume 30, Issue 3

In this paper, a new fatigue reliability analysis method based on saddlepoint approximation (SPA) was proposed for calculating the probability of failure of turbine disk alloy in a low cycle fatigue (LCF) regime. Firstly, two LCF life prediction models based on total strain energy density and Support Vector Regression (SVR) metamodel are presented for turbine disk alloy GH4133 under different loading conditions at 250 °C. Compared with the SWT model, modified Walker model and Response Surface (RS) model, the predicted lives by the proposed models are within a factor of ±2 and a factor of ±1.1 respectively. Secondly, based on the fatigue design criteria, the probabilities of failure are calculated using SPA for the explicit and implicit performance functions using two proposed LCF models and viscosity-based model. These three models have provided the reliability design rules for GH4133. Finally, the failure probabilities curves between SPA and the designed fatigue lives are achieved. The reliability analysis results were found to be in good agreement with the calculated results of test data. These results show that SPA is very apt for the fatigue reliability analysis of turbine disk under different loading conditions using only a small number of samples without any distribution assumptions for random variables. Moreover, it can be used to estimate the system's probability of failure with a large number of random variables or high nonlinearity of performance functions. The effectiveness and accuracy of the combination of the fatigue models and SPA for fatigue reliability analysis are verified using three examples.

Subscale wind tunnel test of an aircraft vehicle is performed at different Mach number, mass-flow and angle of attack. CFD model, corrected by test results, is also presented to predict inlet performance and total pressure distortion. The result shows total pressure recovery decreases and distortion level rises when Mach number increases from subsonic to supersonic speed, AOA is negative and mass-flow value is too large or too small. Compared linear interpolation based on test result of discrete probes, numerical simulation has advantages in showing inlet flow field predicting actual surface distortion level in AIP plane. Swirl distortion is induced by vortex near the fuselage and adjustable ramp and can strengthen total pressure distortion in AIP at negative AOA. And appropriate suction mass-flow coefficient (1.7% to 3%) is beneficial for inlet performance and total pressure distortion control.

The current techniques for investigating the erosion of turbo machineries rely on visual inspections trough boroscopy. However this implies shutting down the power plant in order to make the assessment which leads to operational costs and difficulties. This paper aims to provide a method for monitoring the erosion state of a bladed power plant operated in dusty environments such as the desert by measuring the changes in its acoustic spectrum. The method used for this study is numerical and the findings suggest that there are significant modifications to both the flow field and the acoustic parameters as the blade gets progressively eroded. This paves the way for the development of non-invasive permanent real time diagnostics for turbine engines and power plants.

The research for fuel consumption and pollution reduction in new generation aero engines has indicated intercooling and regeneration as very effective methods for this purpose. Hence, different countries have joined their efforts in common research programs, to develop new gas turbine engines able to reduce considerably the fuel consumption and the ambient impact by means of these two techniques. To study their effects on the engine performance and characteristics, a thermodynamic numerical program that simulates the behavior of a turboprop engine with intercooling and regeneration in different operating conditions has been developed. After the parametric study, and the definition of the design conditions, the off-design analysis is carried on, comparing the main characteristics of the intercooled-regenerated turboprop with those of a conventional engine. Then, once a particular mission profile was fixed, the engine performance, in particular the equivalent power, the fuel consumption and the heat exchanger weight were discussed.

This work investigated features of the deflagration to detonation transition in a curved tube. A number of experiments were performed to acquire the transition rule of DDT, which would provide the design data and theoretical basis for the curved detonation chamber. The content of research is as follows: (1) Flow resistance experiments of nine detonation chambers have been explored. The results show that the spiral configuration can reduce the axial length of DC, and the total pressure recovery coefficient increases with the spiral pitch. (2) Single-cycle detonation experiments have been conducted using the 9 tubes in the resistance experiments. Liquid-gasoline/air is used as the detonative mixture in all the experiments. The detonation experimental results indicate that there is no detonation wave formed in the straight tube, but in all the selected spiral tubes fully-developed detonation waves have been obtained; compared to the straight tube case, the DDT time decrease with the decreasing of the radius of curvature ( RC ) by 6.2%∼19.8% in the spiral detonation tubes.

Partially filled multi-cycle operation usually happens in pulse detonation rocket engines (PDREs). Thrust estimates of PDREs under such operating conditions will provide significant reference for relevant studies. An analytical model for thrust estimates is proposed in the present study. This model is based on Wintenberger' model for impulse calculation and empirical formula for specific impulse prediction of partially filled tubes; moreover, it takes into account performance penalty created by obstacles preliminarily, which are introduced to accelerate deflagration to detonation transition. Four specific analytical models are developed according to three previously proposed empirical formulas and a fitting one proposed for partial filling effect. Comparisons between model predictions and experimental measurements are carried out to validate the reliability of the models. It is found that the models make reasonable estimates and one of the models performs very well over a wide range of conditions. Discussion and analysis on the proposed model and partial filling effect are also performed in this study. Although there are many other factors that should be considered in evaluating thrust of partially filled multi-cycle PDREs, the present study supplies a rapid and effective means for thrust estimation and provides some modeling ideas meanwhile.

The first stage of a transonic multi-stage mixed-flow compressor is numerically studied with and without casing treatment. First, overall performance and behavior of the tip leakage flow are investigated in the compressor stage with a smooth casing. Performance problems emerge in terms of narrow operating range and mismatching between rotor and stator. By investigating the flow field near tip region in detail, remarkable low momentum regions are observed at front and rear portion of rotor passage near suction surface, which grow up and extend further upstream along with mass flow reduced to near stall point, which contributes to a narrow stall margin. Detailed analysis indicates that tip leakage vortex and two types of leakage flow with different flow behaviors are main causes of low momentum regions. Shock-leakage flow interaction and vortex breakdown play an important role in formation of these regions. To improve the flow field near the rotor tip and extend the stable operating range, a circumferential groove casing treatment is then adopted, starting from 10% and ending at 94% axial chord. As expected, the low momentum regions are largely diminished by the casing treatment and stall margin of the investigated compressor stage is effectively enhanced. Fluid from circumferential grooves is injected into blade passage near the suction surface and re-energizes the leakage flow, which makes mainly contribution to manipulation of tip leakage flow and stall margin improvement. Since the pressure difference across blade tip section has a great impact on effectiveness of circumferential grooves, the positions of shock wave and tip leakage flow as well as where the interaction takes place ought to be taken into account through the casing treatment design procedure.

The relatively constant flow of the ocean currents carries large amounts of energy that can be captured and converted to a usable form. In this paper, design and performance simulation of a horizontal axial ocean current turbine (OCT) are studied. Also, study on varying the turbine duct housing design parameters and a series of simulation analyses were performed. Our model simulations with and without the turbine duct housing design were compared with existing literature results and very good agreements were obtained. The results also showed that unducted ocean current turbine power output is proportional to the cube of the current speed. Therefore, if we can increase the flow rate, the ocean current turbine generator performance is expected to be greatly enhanced. By adding the turbine duct housing, under the current velocity of 1 and 2 m/s, parametric studies including the duct housing type, duct entrance area, tilt angle, and duct length-to-diameter ratio are performed. After obtaining the optimized set of design values for our current design, a complete system analysis of the ocean current turbine generator was performed to provide future design guidelines.

In the next few decades, the development of hypersonic aircrafts will be the priority of national defense strategy of many countries, and the significance of the propulsion system applied on the aircraft is self-evident. Hypersonic aircrafts require propulsion devices that can be used independently in a wide range of Mach numbers and flight envelop, however, there have been no such propulsion systems up to now. The multi-mode combined detonation engine, which consists of four modes, is a novel propulsion concept proposed at the beginning of this century for potential application on hypersonic cruise or space access. The multi-mode combined detonation engine is a promising propulsion system which can be used independently in a wide range of Mach numbers and flight envelop. Recently, many countries attach much importance to the research of multi-mode combined detonation engine and great progress has been achieved. This paper attempts to provide an overview of the recent advances in the multi-mode combined detonation engine, and the major progress for each mode is detailed introduced. Nevertheless, there is still a need to pay more attention to the existing problems in the future.