Volume 33, Issue 2 -

Aiming to improve the predictive ability of Walker model for fatigue life prediction and taking the turbine disc alloy GH4133 as the application example, this paper investigates a new approach for probabilistic fatigue life prediction when considering parameter uncertainty inherent in the life prediction model. Firstly, experimental data are used to update the model parameters using Bayes’ theorem, so as to obtain the posterior probability distribution functions of two parameters of the Walker model, as well to achieve the probabilistic life prediction model for turbine disc. During the updating process, Markov Chain Monte Carlo (MCMC) technique is used to generate samples of the given distribution and estimating the parameters distinctly. After that, the turbine disc life is predicted using the probabilistic Walker model based on Monte Carlo simulation technique. The experimental results indicate that: (1) after using the small sample test data obtained from turbine disc, parameter uncertainty of the Walker model can be quantified and the corresponding probabilistic model for fatigue life prediction can be established using Bayes’ theorem; (2) there exists obvious dispersion of life data for turbine disc when predicting fatigue life in practical engineering application.

In this work, a hydrogen-fueled dual-mode scramjet combustor was investigated experimentally. Clean and dry air was supplied to the combustor through a Mach 2 nozzle with a total temperature of 800 K and a total pressure of 800 kPa. The high enthalpy air was provided by an electricity resistance heater. Room temperature hydrogen was injected with sonic speed from injector orifices vertically, and downstream the injector a tandem cavity flame holder was mounted. Except wall pressure profiles, velocity and temperature profiles in and at exit of the combustor were also measured using hydroxyl tagging velocimetry (HTV) and tunable diode laser absorption spectroscopy (TDLAS), respectively. Results showed that combustion occurred mainly at the bottom side of the combustor. And there were also an extreme disparity of the velocity and temperature profiles along the Y -direction, i.e. the transverse direction.

Adjoint method is an important tool for design refinement of multistage compressors. However, the radial static pressure distribution deviates during the optimization procedure and deteriorates the overall performance, producing final designs that are not well suited for realistic engineering applications. In previous development work on multistage turbomachinery blade optimization using adjoint method and thin shear-layer N-S equations, the entropy production is selected as the objective function with given mass flow rate and total pressure ratio as imposed constraints. The radial static pressure distribution at the interfaces between rows is introduced as a new constraint in the present paper. The approach is applied to the redesign of a five-stage axial compressor, and the results obtained with and without the constraint on the radial static pressure distribution at the interfaces between rows are discussed in detail. The results show that the redesign without the radial static pressure distribution constraint (RSPDC) gives an optimal solution that shows deviations on radial static pressure distribution, especially at rotor exit tip region. On the other hand, the redesign with the RSPDC successfully keeps the radial static pressure distribution at the interfaces between rows and make sure that the optimization results are applicable in a practical engineering design.

In this paper, flow separation of a convergent–divergent (C-D) nozzle is placed downstream of a supersonic flow delivered from Mach 2.0 nozzle is investigated. Static pressure measurements are conducted using pressure taps. The flow characteristics of straight and slanted entry C-D nozzle are investigated for various NPR of Mach 2.0 nozzle. The effect of asymmetry at inlet by providing 15°, 30°, 45° and 57° cut is analyzed. Particular attention is given to the location of the shock within the divergent section of the test nozzle. This location is examined as a function both NPR of Mach 2.0 nozzle and test nozzle inlet angle. Some of the measurements are favorably compared to previously developed theory. A Mach number ratio of 0.81 across the flow separation region was obtained.

This study extends the discontinuous Galerkin (DG) methods to simulations of thermoelasticity. A thermoelastic formulation of interior penalty DG (IP-DG) method is presented and aspects of the numerical implementation are discussed in matrix form. The content related to thermal expansion effects is illustrated explicitly in the discretized equation system. The feasibility of the method for general thermoelastic simulations is validated through typical test cases, including tackling stress discontinuities caused by jumps of thermal expansive properties and controlling accompanied non-physical oscillations through adjusting the magnitude of IP term. The developed simulation platform upon the method is applied to the engineering analysis of thermoelastic performance for a turbine vane and a series of vanes with various types of simplified thermal barrier coating (TBC) systems. This analysis demonstrates that while TBC properties on heat conduction are generally the major consideration for protecting the alloy base vanes, the mechanical properties may have more significant effects on protections of coatings themselves. Changing characteristics of normal tractions on TBC/base interface, closely related to the occurrence of coating failures, over diverse components distributions along TBC thickness of the functional graded materials are summarized and analysed, illustrating the opposite tendencies in situations with different thermal-stress-free temperatures for coatings.

To stay relevantly supported , jet-engine researchers, designers and operators should follow changing uses of small and large jet engines, especially those anticipated to be used by/in the next generation, JET-ENGINE-STEERED (“ JES ”) fleets of jet drones but fewer, JES-Stealth-Fighter/Strike Aircraft. [1–19, Figure 5 on WIN-WIN VOLL-ROLL-TARGETING & APPENDIX ]. In addition, some diminishing returns from isolated, non-integrating, jet-engine component studies, vs. relevant, supersonic, shock waves control in fluidic -JES-side-effects on compressor stall dynamics within Integrated Propulsion Flight Control (“ IPFC ”), and/or mechanical JES, constitute key relevant methods that currently move to China, India, South Korea and Japan . [4, 5, 20–31]. The central roles of the jet engine as primary or backup flight controller also constitute key relevant issues, especially under post stall conditions involving induced engine-stress while participating in crash prevention or minimal path-time maneuvers to target. [32–45]. And when proper instructors are absent, self-study of the JES-STVS REVOLUTION is an updating must, where STVS stands for wing-engine-airframe-integrated, embedded stealthy-jet-engine-inlets, restructured engines inside Stealth, Tailless, canard-less, Thrust Vectoring IFPC Systems. [4, 5]. Anti-terror and Airliners Super-Flight-Safety are anticipated to overcome US legislation red-tape that obstructs JES-add-on-emergency-kits-use. [1, 2, 13, 16, 17, 19, 32, 33, 36, 39–41, 43–45].

This paper deals with failure analysis of a low-pressure turbine blade of a straight flow turbojet engine. The blade is made of a wrought precipitation hardened Nickel base superalloy with oxidation-resistant diffusion aluminizing coating. The failure mode is found to be fatigue with multiple cracks inside the blade having crack origin at metal carbides. In addition to the damage in the coating, carbide banding has been observed in few blades. Carbide banding may be defined as inclusions in the form of highly elongated along deformation direction. The size, shape and banding of carbides and their location critically affect the failure of blades. Carbon content needs to be optimized to reduce interdendritic segregation and thereby provide improved fatigue and stress rupture life. Hence, optimization of size, shape and distribution of carbides in the billet and forging parameters during manufacturing of blade play a vital role to eliminate/reduce extent of banding. Reference micrographs as acceptance criteria are essential for evaluation of raw material and blade. There is a need to define the acceptance criteria for carbide bandings and introduce more sensitive ultrasonic check during billet and on finished blade inspection.

Experiment and numerical simulation of flow transition in a compressor cascade with positively curved blade is carried out in a low speed. In the experimental investigation, the outlet aerodynamic parameters are measured using a five-hole aerodynamic probe, and an ink-trace flow visualization is applied to the cascade surface. The effects of transition flow on the boundary layer development, three-dimensional flow separation and aerodynamic performance are studied. The feasibility of a commercial computational fluid dynamic code is validated and the numerical results show a good agreement with experimental data. The blade-positive curving intensifies the radial force from the endwalls to the mid-span near the suction surface, which leads to the smaller scope of the intermittent region, the lesser extents of turbulence intensity and the shorter radial height of the separation bubble near the endwalls, but has little influence on the flow near the mid-span. The large passage vortex is divided into two smaller shedding vortexes under the impact of the radial pressure gradient due to the positively curved blade. The new concentrated shedding vortex results in an increase in the turbulence intensity and secondary flow loss of the corresponding region.

An experimental study of the recirculation zone and mixing lengths for bluff-body stabilized flames is conducted at non-reactive conditions. This paper reports the prediction of recirculation zone length from dynamic pressure measurements. The auxiliary turbulence created from the wall of the combustor is also studied and maintained to levels as low as 5%. The experiments are conducted by varying the velocity from 5 m/s to 8 m/s for V-Gutters bluff-body with induced angles of 60, 90 and 120 o . These gutters are maintained at same blockage ratio so that gutter angle to flow velocity is studied. It is inferred from the experiment that as the velocity in the duct increases, the length of the recirculation zone varies 5 mm for all V-Gutter angle. However, an increase in the V-gutter angle is observed to greater effect than an increase in the velocity, recirculation zone length which varied from 70 mm for 60 o V-gutter to 150 mm for 120 o V-gutter. Simultaneously a sharp reduction in shear distribution along the length of the combustor are observed, it influences in understanding the mixing characteristics in combustion.

Fan discs are critical components of an aero engine. In this paper, contact stress and life prediction of a turbine fan disc were investigated. A simplified pin/disc model was conducted to simulate the practical working condition under applied loads using finite element (FE) analysis. This study is devoted to examining the effects of interface condition of pin/disc such as gap and coefficient upon the maximum stress. The FE model indicated that the maximum stress occurs at the top right corner in the second pin hole, and larger gap or friction coefficient has a significant effect on the maximum stress. In addition, FE analysis without considering friction is also conducted. The results show that the dangerous point is similar to the result which considers friction and the stress state is relatively larger than that of considering friction. Finally, based on FE analysis result, life prediction for the fan disc is conducted to combine the material S-N curve, mean stress effects and concentration stress factor obtained by means of FE method.

In order to obtain better integrated performance of aero-engine during the conceptual design stage, multiple disciplines such as aerodynamics, structure, weight, and aircraft mission are required. Unfortunately, the couplings between these disciplines make it difficult to model or solve by conventional method. MDO (Multidisciplinary Design Optimization) methodology which can well deal with couplings of disciplines is considered to solve this coupled problem. Approximation method, optimization method, coordination method, and modeling method for MDO framework are deeply analyzed. For obtaining the more efficient MDO framework, an improved CSSO (Concurrent Subspace Optimization) strategy which is based on DOE (Design Of Experiment) and RSM (Response Surface Model) methods is proposed in this paper; and an improved DE (Differential Evolution) algorithm is recommended to solve the system-level and discipline-level optimization problems in MDO framework. The improved CSSO strategy and DE algorithm are evaluated by utilizing the numerical test problem. The result shows that the efficiency of improved methods proposed by this paper is significantly increased. The coupled problem of VCE (Variable Cycle Engine) conceptual design is solved by utilizing improved CSSO strategy, and the design parameter given by improved CSSO strategy is better than the original one. The integrated performance of VCE is significantly improved.