Band 40, Heft 4

Design and aerodynamic performance analysis of a variable geometry axisymmetric inlet was carried out for tandem scheme turbine-based combined cycle (TBCC) propulsion system. The operation Mach number of the inlet was between 0 and 4. The design point was chosen as Mach number 4.0 in this paper. The determination of external and internal compression and the design method of annular to circle diffuser were illustrated. The inlet was unstart under Ma 3.0 without adjustment. Then, a variable scheme was designed to ensure self-start of the inlet and match the requirement of mass flow rate during the whole flight envelope. And four supports were used to fix the spike. According to the 3D numerical simulations, the total pressure recovery was 0.52 at Ma 4.0 at critical condition and the mass flow rate was consistent with the requirement at different flight Mach number.

In the present work, a three-dimensional numerical analysis of multi-hole film cooling on an adiabatic flat plate with combined backward and forward injection holes is carried out at varying velocity ratios from 0.5 to 2.5. The effect of three different multi-hole arrangements, square-diamond, long-diamond, and super-long-diamond with constant perforated percentage (3.27%), on film cooling performance is studied. Combining backward and forward injection holes help in achieving higher velocity boundary layer thickness in the multi-hole arrangement. It is observed that super-long-diamond arrangement provides the highest film cooling effectiveness than square-diamond and long-diamond arrangements and favours early development of coolant film layer.

There is inevitably a performance deviation between an engine model and an actual engine that is influenced by unpredictable factors such as the unsuspected environmental conditions and the natural performance degradation in the process of use. Because the engine model precision largely depends on the accuracies of the component maps, it is possible to revise the engine model to determine a better trend for the engine performance from recorded measurements by adjusting the maps. This paper presents a new method for updating the variable geometry component maps of a variable cycle engine (VCE) by using a set of scaling factors estimated with the cubature Kalman filter (CKF). A mapping function is created between the scaling factors and the component characteristic scaling coefficients for the adjustments of the maps. The proposed method is applied to a VCE model according to the VCE benchmark steady-state performance data. The results show that the maximum simulation error of the engine steady-state model decreases from 5.33 to 0.93%, and the CKF-based adaptation method provides a much faster computing rate than the particle swarm optimization (PSO) based adaptation method, which verifies the effectiveness and engineering applicability of the variable geometry characteristic adaptive correction method.

In this work, a circumferential groove on the shroud was applied to a centrifugal compressor and optimized to improve the stability of the compressor. To reduce the number of design variables before optimization, the position of the casing groove that maximizes the stall margin of the compressor was firstly selected. The width and height of the casing groove were optimized using the RSM (Response Surface Method) and the full factorial design of experiments to increase the stall margin further. It was observed that the optimized case with the casing groove can increase the stall margin by 7.0% but the adiabatic efficiency at the design condition decreases by 0.99%, in comparison to the reference case without the groove. Finally, the flow field of both cases with and without the optimized casing groove was compared to each other to analyze effects of the casing groove on the stall margin and design efficiency. The casing groove decreases the leakage flow across the impeller tip and weakens the tip leakage vortex, consequently decreasing the blockage near the shroud and improving the stall margin. However, it was found that the spillage of the flow from the casing groove increases the mixing loss near the shroud.

Significant increase in task complexity for modern gas-turbine propulsion systems drives the need for future advanced cycles’ development. Further performance improvement can be achieved by increasing the number of engine controls. However, there is a lack of cycle analysis tools, suitable for the increased complexity of such engines. Towards bridging this gap, this work focuses on the computation time optimization of various mathematical approaches that could be implemented in future cycle-solving algorithms. At first, engine model is described as a set of engine variables and error functions, and is solved as an optimization problem. Then, the framework is updated to use advanced root-finding paradigms. Starting with Newton-Raphson, the model is improved by applying Broyden’s and Miller’s schemes and implementing solution existence validation. Finally, algorithms are compared in representative condition using increasingly complex turbojet and adaptive cycle turbofan configurations. As evaluation cases become more time consuming, associated time benefits also improve.

This paper presents the steady numerical investigation on SW/TLV interaction with SST turbulence model at two characteristic operating conditions for a transonic fan rotor, NASA Rotor 67. The main purpose of the present work is to reveal the main flow structures and properties during the SW/TLV interaction, and a theoretical criterion for vortex stability is engineeringly utilized to determine such shock wave-induced vortex stability. The validations for all numerical schemes have been conducted by comparing the RANS solutions with detailed experimental data before the analyses of flow phenomenon and mechanism. The simulation results indicate that numerical methods used in NUAA-Turbo 2.0 solver, independently developed by our team, enable to accurately capture the complex flow structures including shock wave and vortex systems within the blade passages, especially in the tip region. Similar to wing-tip vortex created by vortex generator, the TLV has the same wake-type characteristics. The flow pattern generated by such interaction is characterized by the bulged-forward shock front followed by a subsonic flow region and a slight expansion of vortex core. No apparent vortex breakdown was examined by both intuitive visualization of three-dimensional vortex structure and a theoretical criterion.

As a novel structural damper, the unique structural characteristics of the integral squeeze film damper (ISFD) solve the nonlinear problem of the traditional squeeze film damper (SFD), and it has good linear damping characteristics. In this research, the experimental studies of ISFD vibration reduction performance are carried out for various working conditions of unbalanced rotors. Two ball bearing-rotor system test rigs are built based on ISFD: a rigid rotor test rig and a flexible rotor test rig. When the rotational speed of rigid rotor is 1500 rpm, ISFD can reduce the amplitude of the rotor by 41.79%. Under different unbalance conditions, ISFD can effectively improve the different degrees of unbalanced faults in the rotor system, reduce the amplitude by 43.21%, and reduce the sensitivity of the rotor to unbalance. Under different rotational speed conditions, ISFD can effectively suppress the unbalanced vibration of rigid rotor, and the amplitude can be reduced by 53.51%. In the experiment of the unbalanced response of the flexible rotor, it is found that ISFD can improve the damping of the rotor system, effectively suppress the resonance of the rotor at the critical speed, and the amplitude at the first-order critical speed can be reduced by 31.72%.

This study investigates the turbulent flow field downstream of V-gutters using unsteady numerical modelling. An important domain of application of the vortex shedding induced by the V-gutters is the flame stabilization in high speed combustion systems which find extensive applications in aerospace engineering. In view of this, the present study analyses the impact of the V-gutter geometry, as characterized by the included angle, on inducing vortex motion in the wake. Transient simulations are carried out for three values of the semi-span angle, α  = 30°, 45° and 60°. Based on the analysis of the saddle point and the vortex shedding frequency, the study shows that an increase in span angle within this range, favours the effectiveness of the method in flame stabilization. Though the simulations are done for cold flow, the dominant mechanism of vortex shedding is adequately addressed in the analysis.

The design of a two-dimensional variable geometry inlet which applied to a tandem type turbine-based combined cycle (TBCC) propulsion system was investigated in the present paper through three-dimensional simulations and wind tunnel tests. The operation Mach number range was between 0 and 3. A multi-ramp geometry scheme was adopted to achieve acceptable performance at different inflow Mach number. The first ramp angle was fixed whilst the angles of the second and the third ramps were variable at different inflow Mach numbers. The Mach numbers at throat region were maintained between 1.3 and 1.5 at different inflow Mach numbers according to this variable geometry scheme. A fixed geometry rectangular-to-circular shape diffuser was adopted to improve aerodynamic performance of the inlet. Three-dimensional numerical simulations were carried out between Ma 1.5 and Ma 3.0. The results indicated that good aerodynamic performance can be achieved at different inflow speed. At the design point, total pressure recovery of the inlet was 0.66 at critical condition. Wind tunnel validation experiment tests were conducted at Ma 2.0, showing the movement of terminal shock wave from downstream to upstream as the back pressure increased. The inlet operated at supercritical, critical and subsonic conditions at different back pressure.

This paper presents the computational study carried out on an aero gas turbine combustor to assess important performance parameters. The CFD results are compared with experimental dataobtained from the full scale combustor tested at ground test stand simulating various operational conditions. The CFD predictions have agreed very well with the experimental data. The model is then extended to predict combustor exit temperature pattern factors, pressure loss, and combustion efficiency and exhaust gas constituents over a wide range of operating pressure and temperature conditions. The paper also presents the studies carried out on the effect of atomizer spray cone angle, particle size and fuel flow variations expected due to manufacturing tolerances in various flow passages as well as due to operational degradations on temperature pattern factors. The pattern factors are also analyzed on cold and hot day environment. The radial pattern factor (RPF) at mid height is found to increase as altitude increases from sea level to 12 km. Spray cone angle is found to have a predominant effect on temperature non-uniformity at exit, lower cone angle increasing both radial and circumferential pattern factors. The findings of this study are valuable inputs for engine performance estimation.

The lean blowout (LBO) limit is crucial for gas turbine combustor in the aero engine. The effect of atomization of liquid fuels on the LBO limit is needed to be further studied. On the other hand, the effects of atomization on the LBO limit can be neglected if gas fuels are utilized in a combustor. Thus, the comparative experiment between liquid fuel and gas fuel can be utilized to study the effects of atomization performance of liquid fuels on the LBO limit. In this paper, the LBO limit for a gas turbine combustor utilizing methane is studied experimentally. Seven kinds of combustor configurations are chosen for the experimental test. The LBO limits are obtained for all the chosen combustors. The variation of the LBO limit with the combustor configuration for both methane and aviation kerosene exhibits the similar tendency, i.e., the LBO limits utilizing methane are slightly larger than those utilizing aviation kerosene for the same combustor. Further, the atomization performance has little effects on the LBO limits for the present combustor configurations at the present operating conditions where the SMD for the fuel atomizer utilizing aviation kerosene is about 10 μm.

In this study, the authors propose an optimization process to design the baseline rotor of a supersonic through-flow fan (STFF) at an inlet Mach number of 2.0 based on Genetic Algorithm. Unlike the improvement in performance brought about by the pre-compression of conventional supersonic profiles in the presence of axial pressure flow, pre-compression did not help improve the performance of the rotor of the STFF. The efficiency of elements of the blade at spanwise heights of 10 %, 50 %, and 90 % increased by 2.47 %, 1.95 %, and 2.49 %, respectively. The performance of the rotor of the STFF that was reconstructed by stacking the optimized elements of the blade was improved at the design point as well as in off-design conditions by using three-dimensional computational fluid dynamics (CFD) simulations. The performance of the blade also improved considerably, with increases of by 2.46 % and 9.59 % in its isentropic efficiency and the overall pressure ratio, respectively.