Volume 31, Issue 3

Advanced jet engines do not operate in an application vacuum. Their optimal use in advanced military applications drives much of their basic innovative research and development, especially when new needs arise in the rapidly changing domains of stealth-agile, fighter aircraft and tailless-stealth, Jet-Steered, Unmanned Air Vehicles (JS-UAV). For these reasons we periodically update this Journal with new trends that affect, and sometimes control, research and development of future jet-engines. One relevant example is the recently unmasked RQ-180 stealth-tailless drone, which is an improved version of the smaller, RQ-170 captured by Iran. Most important, with the new X-47B/C tailless-stealth JS-UAV, it is to dominate future uses of fuel-efficient jet-engines, especially for operating in dusty environments. The RQ-180 has been secretly designed and funded since 2008. It is based on a classified, 1986, parent Israeli Patents 78402, which protect hundreds design and testing trade secrets taken from 1986 to 1997 by the United States Government (USG) via classified contracts with USG-Contractors Boeing, Lockheed, General Dynamics and General Electric, as revealed by a December 6, 2013 Aviation Week [1–3] and U.S. Court of Appeals for the Federal Circuit, Case 2014-5028, Docket 12 [4]. The new RQ-180 design explains the recent U.S. Air Force ISR shift away from “permissive” environments – such as Iraq and Afghanistan, where non-stealthy Global Hawk and General Atomics' Reaper operate – toward new missions in highly “contested” or strongly “denied” enemy airspaces.

This paper proposes a model for characterizing the flow pattern of a centrifugal impeller attending to the severity of the reverse flow. The model assumes 1) a definition of an escaping particle as the one that flows in every operational point from the trailing edge towards the leading edge of the impeller blades, and 2) a characterization of flow where an operational point is said to have a theoretical flow pattern if it is not possible to establish a fully-reversed escaping particle on it. Therefore, the first part of the article is focused on defining an escaping particle for a centrifugal compressor. The model locates over the map of a centrifugal impeller the line that splits the map in two regions: the region on the right hand side, where a theoretical flow pattern can exist, and the region on the left, where a theoretical flow pattern cannot exist. Therefore, the locus of this line marks a frontier where the expected performance of the impeller cannot be sustained as high as expected. The second part of the article uses a high-performance commercial centrifugal impeller wheel for contrasting the model. A qualitative characterization of the surge line, conclusions and discussions are presented.

This paper presents the numerical investigation through time marching technique for the flow through a contour nozzle kept in a supersonic stream. In such a flow a detached shock would appear before the inlet of the nozzle, the shock stand-off distance is determined for various freestream conditions. Starting from arbitrarily assumed initial conditions, the steady state solution for the flow field, shock wave shape and location is achieved to second order accuracy. The result shows that the shock stand-off distance is influenced by upstream Mach numbers. The shock stand-off distance is compared with commercial software code and satisfactory results are obtained. The flow properties along the centerline obtained from the computation are found to be in good agreement with the simple one-dimensional solution. Even for coarser meshes, the shock appearing in the divergent section of the nozzle captured by the present calculation agrees well with one-dimensional normal shock solution.

The paper consists in modelling and simulation of the combustion in a turbojet engine in order to find optimal characteristics of the burning process and the optimal shape of combustion chambers. The main focus of this paper is to find a new configuration of the aircraft engine combustion chambers, namely an engine with two main combustion chambers, one on the same position like in classical configuration, between compressor and turbine and the other, placed behind the turbine but not performing the role of the afterburning. This constructive solution could allow a lower engine rotational speed, a lower temperature in front of the first stage of the turbine and the possibility to increase the turbine pressure ratio by extracting the flow stream after turbine in the inner nozzle. Also, a higher thermodynamic cycle efficiency and thrust in comparison to traditional constant-pressure combustion gas turbine engines could be obtained.

Four different secondary airflow angles for the turbine inter-guide-vane burners with trapped vortex cavity were designed. Comparative analysis between combustion performances influenced by the variation of secondary airflow angle was carried out by using numerical simulation method. The turbulence was modeled using the Scale-Adaptive Simulation (SAS) turbulence model. Four cases with different secondary jet-flow angles (−45°, 0°, 30°, 60°) were studied. It was observed that the case with secondary jet-flows at 60° angle directed upwards (1) has good mixing effect; (2) mixing effect is the best although the flow field distributions inside both of the cavity and the main flow passage for the four models are very similar; (3) has complete combustion and symmetric temperature distribution on the exit section of guide vane (X = 70 mm), with uniform temperature distribution, less temperature gradient, and shrank local high temperature regions in the notch located on the guide vane.

The fuel control method of the liquid kerosenefueled scramjet combustor was investigated experimentally in this article. The primary goal was to experimentally demonstrate the possibility of using the closed-loop control with a combustor wall pressure feedback in the scramjet combustor. The open-loop control and the closed-loop control experiment were investigated by using the direct-connected test facility of Northwestern Polytechnical University. The typical flow conditions at the combustor entrance were: Mach number Ma = 2.0, totaltemperature Tt ≈ 880 K and total-pressure Pt = 700 ~ 800 kPa, which corresponded to a flight condition with Mach number 4.0. The monotone relationship between the wall pressure and the engine thrust was experimentally demonstrated by the open-loop control experiments. The closed-loop control experiments demonstrated that it was feasible to use wall pressure or pressure ratio as the feedback parameter in the closed-loop control, and the results indicate that it can precise control the fuel applying and shows a good performance.

Gas turbine engine gas path fault diagnosis is closely related technology that assists operators in managing the engine units. However, the performance gradual degradation is inevitable due to the usage, and it result in the model mismatch and then misdiagnosis by the popular model-based approach. In this paper, an on-line integrated architecture based on nonlinear model is developed for gas turbine engine anomaly detection and fault diagnosis over the course of the engine's life. These two engine models have different performance parameter update rate. One is the nonlinear real-time adaptive performance model with the spherical square-root unscented Kalman filter (SSR-UKF) producing performance estimates, and the other is a nonlinear baseline model for the measurement estimates. The fault detection and diagnosis logic is designed to discriminate sensor fault and component fault. This integration architecture is not only aware of long-term engine health degradation but also effective to detect gas path performance anomaly shifts while the engine continues to degrade. Compared to the existing architecture, the proposed approach has its benefit investigated in the experiment and analysis.