Volume 29, Issue 4

The zone of influence and the positive pressure field over the nose, for a blunt-nosed body, with and without spike, were measured by direct means. Flow past the body, with and without spikes, was visualized in a two-dimensional water channel, at a specific Reynolds number. It is found that, the spike reduces the radius of curvature of the approaching streamline, and the deflection of the streamline is towards the shoulder of the basic body, resulting in a narrow zone of the positive pressure hill at the nose. The spike with hemispherical nose reduces the zone of influence drastically from 1.75 to less than 1.60. The pressure hill height for hemispherical nose is found to be smaller, compared to other shapes of the spike nose studied. The hill heights for the flat faced spike with cylindrical and square stems are comparable. It is found that, the size of the vortex at the junction of the spike and basic body is the largest for the hemispherical nose.

This study numerically investigated the flow return system performances of two-stage high pressure fans. The parameters include inflow angle, length, exit width of vaneless diffuser and the blade number of deswirl vane. This study used the tested impeller to verify the accuracy of numerical calculations, with good consistency. The analytic results showed that the overall pressure loss of impeller and flow return system is effectively reduced when the inflow angle of vaneless diffuser is 15° and tangential to wheel velocity ratio is 0.46. A vaneless diffuser was required and its length referred to the maximum outer diameter of the volute casing. The performance of a contracted-wall vaneless diffuser was not improved over that of a parallel-wall one. The suggested number of blades in a deswirl vane was 12. The downstream affects the upstream in several areas inside the flow return system, thus, the flow return system must be calculated integrally.

Studies on the cooling effectiveness distribution between scalelike laminated plate cooling duct and routine round film hole with Numerical Investigation show that the inner duct structure of the scalelike laminated plate can even the temperature and velocity on the exit of the film hole and the arrangement of film holes can even the coverage of the film on the wall surface to lower the decline rate of the wall surface temperature; The scalelike inner duct structure can cause turbulence in the cooling air when they meet at the intersection, hence benefit the convection exchange inside the mental surface; The scalelike laminated plate can also twice the cooling effectiveness of the routine round film hole.

Recently, a theoretical criterion to calculate the stability of an axial-flow compressor rotor has been presented in the scientific literature. This theoretical criterion was used for determining the locus of the stability line over the rotor map and for predicting the post-stall evolution of the constant-speed line of a rotor. The main objective of this paper is to improve the predictions of such a model. To do that, the paper proposes a different characterization of the characteristic azimuthal length and a calculation of the ratio of specific heats based on a polytropic exponent. Thanks to these new values, the model predicts two bifurcation points in the behaviour of the flow: the inception point of the instability and the surge point. Experimental data from a pure axial compressor are used to validate the model showing that the prediction of the flow coefficient at the surge point has an error inferior to 5%. The paper presents a discussion about how the model can explain the different behaviours exhibited by the same rotor when the flow coefficient is reduced.

Aircraft engine components are subjected to variable amplitude load conditions usually they tend to experience fatigue damage. Many fatigue damage accumulation theories have been put forward to predict the fatigue lives of structure components. Under different load spectra, the aircraft engine has different working life. In this paper, based on the linear damage rule, the fatigue life of an aircraft engine was estimated by considering the load spectrum difference factor. According to one aircraft engine used for different load spectrums, the relationship between the load spectrum and the life is discussed. Moreover, the fatigue life of this engine used in A-type bomber is estimated based on the different load spectrum of the B-type bomber. A good agreement is found between the estimated results and real life data.

The numerical analysis of certain safety related problems presents serious difficulties, since the large number of components present leads to huge finite element models that can only be solved by using large and expensive computers or by making rough approaches to the problem. Tangling, or clashing, in the turbine of a jet engine airplane is an example of such problems. This is caused by the crash and friction between rotor and stator blades in the turbine after an eventual shaft failure. When facing the study of an event through numerical modelling, the accurate simulation of this problem would require the engineer to model all the rotor and stator blades existing in the turbine stage, using a small element size in all pieces. Given that the number of stator and rotor blades is usually around 200, such simulations would require millions of elements. This work presents a new numerical methodology, specifically developed for the accurate modelling of the tangling problem that, depending on the turbine configuration, is able to reduce the number of nodes up to an order of magnitude without losing accuracy. The methodology, which benefits from the cyclic configuration of turbines, is successfully applied to the numerical analysis of a hypothetical tangling event in a turbine, providing valuable data such as the rotating velocity decrease of the turbine, the braking torque and the damage suffered by the blades. The methodology is somewhat general and can be applied to any problem in which damage caused by the interaction between a rotating and static piece is to be analysed.

The paper presents an attempt to determine weather the architecture difference between the two spool and three spool turbofan engines can have any significant effect uppon their use in super circulation wing (SCW) aircraft. Such aircraft have been experimented with since 1970's and have been revived because of growing interest in silent STOL aircraft. The approach used is statistical, i.e. actual data provided by engine manufacturers and regulating authorities such as ICAO and EASA was used to derive certain relevant parameters which were then analyzed and conclusions were formulated.

To improve the understanding of high performance centrifugal compressors found in modern aircraft engines, the aerodynamics through these machines must be experimentally studied. To accurately capture the complex flow phenomena through these devices, research facilities that can accurately simulate these flows are necessary. One such facility has been recently developed, and it is used in this paper to explore the effects of averaging total pressure and total temperature measurements to calculate compressor performance. Different averaging techniques (including area averaging, mass averaging, and work averaging) have been applied to the data. Results show that there is a negligible difference in both the calculated total pressure ratio and efficiency for the different techniques employed. However, the uncertainty in the performance parameters calculated with the different averaging techniques is significantly different, with area averaging providing the least uncertainty.

A small pulse detonation rocket engine (PDRE) was used as a predetonator to initiate detonation in its ejector. The detonation products discharged from the PDRE was not only ignition source for the ejector but also primary flow which entrained air from environment into the ejector. Stoichiometric liquid kerosene and gaseous oxygen were used as reactants for the PDRE. While in the ejector injected liquid kerosene was used as fuel and entrained air was used as oxidizer. Reactants in the ejector were ignited by the detonation wave and products discharged from the PDRE. Detonation was successfully initiation in present experiments. It was found that flame propagation upstream at the entrance of the ejector was inevitable, which affected the detonation initiation process in the ejector. Disks with orifices were placed at the entrance of the ejector to weaken the flame propagation upstream effect, which would affect the air flow entraining process, but the results show it worked.

Fault tree analysis is an important tool for system reliability analysis. Fuzzy fault tree analysis of uncontained events for aero-engine rotor is performed in this article. In addition, a new methodology based on fuzzy set theory is also used in fault tree analysis to quantify the failure probabilities of basic events. The theory of fuzzy fault tree is introduced firstly. Then the fault tree for uncontained events of an aero-engine rotor is established, in which the descending method is used to determine the minimal cut sets. Furthermore, the interval representation and calculation strategy is presented by using the symmetrical L-R type fuzzy number to describe the failure probability, and the resulting fault tree is analyzed quantitatively in the case study.