Band 41, Heft 4

In order to determine the optimal wall pressure location as the controlled parameter for kerosene-fueled scramjet, a combined approach of numerical simulation and experimental analysis was employed. The results indicate that with an increase in the equivalence ratio of kerosene, the combustor exhibits consecutive modes, including pure scramjet mode, dual-mode scramjet mode, dual-mode ramjet mode, and pressure disturbance to the isolator inlet mode. The peak pressure ratios corresponding to the mode transition boundaries are 1.95, 3.26, and 3.57, with respective isolator outlet Mach numbers of 1.62, 1.0, and 0.74. The equivalence ratio corresponding to mode transition increases with an increase in the combustor inlet total temperature. Considering the dynamic characteristics of wall pressures at different locations, the position at x / L  = 0.46 demonstrates the best linearity, highest sensitivity, and greater stability. Hence, it is the most suitable choice as the controlled parameter for fuel flow in this kerosene-fueled scramjet.

Based on the trajectory optimization method of the Gauss pseudospectral, an aircraft/engine matching method is established for the turbine-based combined cycle (TBCC) engine. For a horizontal-takeoff hypersonic aircraft designed at Mach 5, a thrust-matching analysis of the TBCC engine is performed, and a rocket is integrated for further optimization design. The results show that the aircraft for boost missions should adopt the TBCC thrust with a takeoff thrust-to-weight ratio of 99.8 % to reduce the acceleration time and fuel consumption. In contrast, due to the low thrust-to-weight ratio of the TBCC engine, a high-thrust TBCC increases the inert weight in the cruise phase. Therefore, the aircraft designed for cruise missions should adopt the takeoff thrust-to-weight ratio of 92.0 %. Introducing a rocket whose maximum thrust is 10 % of the takeoff weight could assist the aircraft in overcoming the problem of the “thrust pinch” during the transonic and mode transition. With the assistance of rockets, the optimal takeoff thrust-to-weight ratio is 65.3 % for cruise aircraft, and the cruise range is increased by 18 %. While for the boost aircraft, adopting an optimal TBCC of 86.8 % takeoff thrust-to-weight ratio, the introduced rocket could reduce the fuel consumption and the TBCC engine weight by 4 %.

Short take-off/vertical landing (STOVL) engine is an emerging power source for fighter planes that can perform takeoff and landing operations in tight quarters. As the aforementioned function requires specific component design and matching, it is vitally necessary to study the impacts of various component states on engine performance and stability. This paper develops a performance model for STOVL engines based on nonlinear component models and validates the modelʼs correctness. Variable component working conditions and settings are altered, engine performance is forecasted and observed, and the mechanism underlying this trend is investigated. The results indicate that the performance of the STOVL engine is greatly influenced by the design parameters of the gas path, such as the bypass ratio and the injection volume of the roll nozzle, and is highly dependent on the correct adjustment of the rear bypass injector and the throat area of the nozzle.

A double-wall cooling configuration including impingement holes and film holes with conformal pins is studied by CFD numerical simulation in this paper. The influences of three different injection directions of film holes at six blowing ratios ranging from 0.4 to 2.5 on conjugate heat transfer characteristic are investigated. Impingement-only arrangements with corresponding pin directions are adopted to illustrate the film cooling gains and the impingement cooling contribution. The results indicate that the effect of different pin directions on overall cooling effectiveness is non-significant. At low blowing ratio, the forward injection arrangement has higher overall cooling effectiveness since greater gain of film cooling, and is transcended by reverse injection arrangement as coolant supplement increases. Moreover, although the cooling effectiveness for normal injection case is the lowest within the whole range of mass flow rate, it possesses the most uniform distribution of cooling effectiveness in downstream region at high blowing ratio.

To achieve a fast and reliable mode transition of the Ma 0–6 external parallel TBCC engine, a design method for the transition state control law is developed. The design parameters of the TBCC engine are obtained based on the mission requirements of hypersonic aircraft. For the key mode transition process of the combination engine, a segmented mode transition strategy is formulated with the objectives of smooth thrust and continuous flow. The mode transition control law of the combination engine is designed and optimized using the Hooke-Jeeve Direct Search Method. The results demonstrate that the proposed design method for the mode transition control law can effectively enable a fast and smooth mode transition of the combination engine, with a total duration of 7 s. During the mode transition process, the total thrust fluctuation during the transition from the turbofan windmill state to the turbofan shutdown state is the largest, and the maximum relative error between the total thrust and the expected value during the mode transition process is 0.4 %.

The PSC (Performance Seeking Control) based on CPSM (Compact Propulsion System Model) has been verified by NASA. However, the CPSM has poor accuracy at off-design points. Therefore, a new basepoint schedule method is proposed to improve the CPSM accuracy at off-design points. At the off-design point, the thermodynamic parameters which is a function of temperature is an importance factor that influence the accuracy of model based on parameter corrections. Therefore, the temperature of fan inlet is taken into account during scheduling the basepoint vector. The simulations have shown that the accuracy of CPSM is at its best when the engine operates at a point where the temperature of the fan inlet is equal to the one of the basepoint. With the increase or decrease of the temperature of the fan inlet, the modeling errors of CPSM will increase. The simulations also demonstrate that the relative errors of the improved CPSM decrease significantly compared to those of the conventional CPSM at the off-design point.

A coaxial jet that has one application in gas turbine combustion chambers where there is a requirement for very good mixing of air and fuel for complete and stable combustion. A study has been carried out to improve the mixing of annular and central jets in the confined space of the combustor. The combustor model is used with two confined coaxial air-swirling jets under non-combustion conditions. The influence of a heated central jet and the effect of blockage at different locations in confined spaces on mixing has been studied. To understand mixing in jets in confinement, the center line axial velocity profile and streamlines have been presented at different locations. The computational results are obtained using commercial CFD software, ANSYS FLUENT 18.0, with a standard k–ω turbulence model. After validation with the existing literature, parametric studies have been carried out with co-swirl and counter-swirl conditions of annular and central jets at different blockage locations in confinement. The results show a good understanding of the mixing process of two jets in confinement.

In modern applications, demand for the rotor shafts made of FGM has increased in aerospace, gas turbine and turbo jet engines due to its improved dynamic characteristics at elevated temperatures. The dynamic vibration response of a functionally graded Jeffcott rotor-bearing system with induced porosities has been studied for non-uniform porosity distributions. The functionally graded (FG) shaft’s radial direction exhibits continuous variation in the material qualities. The FG rotor’s cross-sectional material properties have been graded using the exponential law. The temperature gradients have been considered to be varied across the cross-section based on exponential temperature distribution. For the first time, using the FE method, it has been possible to study the impact of induced porosities on the free vibration frequencies and steady-state vibration responses of the functionally graded rotor-bearing system. This has led to the development of a two-node porous functionally graded rotor element with uneven porosity distributions using the Timoshenko beam theory. It has been observed that the free vibration frequencies and the FG rotor-bearing system critical speeds are affected by increased porosity and thermal gradients. The frequency responses demonstrate that as temperature gradient and volume fraction of porosity rise, the FG rotor-bearing system’s critical speeds shift to the left.

This paper numerically studied the aerodynamic performance of an intake duct affected by the ground effect on a mobile test bench. The simulations were conducted under no wind and headwind conditions. The time evolution of the ground effect indicates that the coherent structure of the vortex system is mainly composed of the ground vortex, the horse-shoe vortex, and the creeping vortex under no wind condition. And it is mainly composed of the ground vortex, the trailing vortex, and the creeping vortex under headwind condition. Compared to the results under no wind condition, the integral vorticity of the ground vortex is larger than that under the headwind condition. The difference of the total pressure recovery coefficient is small, and the total pressure distortion index is large. The results show that with the decrease of the velocity at the intake duct outlet, the intensity of the ground vortex decreases, and the total pressure recovery coefficient at the intake duct outlet increases.

The process of shock wave focusing can make the strength of shock waves be continuously accumulated and turned into detonation wave in Pulse Detonation Engine (PDE). However, its effective application needs the inlet jets be in high temperature and velocity, which is difficult to be satisfied under certain conditions. Therefore, in this paper, metal magnesium assisted detonation initiation is proposed and the effect of magnesium particle addition on the shock wave focusing process in a kerosene-fueled PDE with cavity configuration is investigated through numerical simulation. The result showed that when the temperature of the premixed fuel/air jets injected in opposite direction was set as 650 K, the collision of leading shock waves on the central axis was the main source of energy deposition and the shock wave focusing could make the detonation be initiated in the cavity. When the temperature of jets is reduced to 550 K, fuel ignition and detonation could not be achieved through shock wave focusing. Then adding metal magnesium particles into the combustor made the energy deposition be enhanced and the detonation be induced. The diffusion of metal particles can significantly change the structure, motion, merging and dissipation of vortices in the flow field. Generally, the shock wave focusing process is basically not affected with metal particles injection. Therefore, this method can be successfully employed for detonation initiation in the cavity when the fuel/air premixed jet temperature is not high for PDE.

Solid Oxide Fuel Cell (SOFC) integrated with a Gas Turbine (GT) is a highly efficient way to convert chemical energy of hydro carbon fuel into electrical energy. SOFC-GT is a high temperature, high pressure system and features a fuel flexibility, which makes it easier to retrofit into the existing power plants. A lot of research had already taken place to make SOFC-GT more fuel efficient and cost effective as well. SOFC-GT is a complex thermodynamic electro-chemical system that has numerous input variables that affect the overall efficiency of the system. In order to further develop the SOFC-GT technology and future research, clear understanding is required to analyze relationship between the input and output parameters. One way to analyze the input/output relationship over the range of parameter variations is to conduct a probabilistic sensitivity analysis. A probabilistic sensitivity analysis of 19 input variables has been conducted to understand how each variable would affect the net power output and overall efficiency of a SOFC-GT system.

Because of its narrow and aerodynamic design, cooling the trailing edge of a turbine blade offers an essential problem. For the trailing edge internal cooling of gas turbine blades, wedge-shaped channels are used. Pin-fin arrays play a critical role in enhancing heat transmission at the trailing edge of turbine blades. The flow and heat transfer properties of a number of wedge channels with staggered pin-fins of varied shapes are used to develop a design with a more effective heat transfer effect and reduced flow resistance. The present study compares the combination of circular pin fins and oblong pin fins in a wedge duct with a Reynolds number range of 12,000–85,000 and a constant heat flux (surface) of 3280 W/m 2 applied to the bottom wall, top wall, and surfaces of the fins. The results indicate that the pressure loss is less in circular fins at the centre and oblong fins side by side, and it is 23.67 % less than circular and oblong fins in opposite positions. The Heat transfer coefficient for circular and oblong fins in opposite positions is 16.72 % higher than circular fins at centre and oblong fin side by side. The thermal performance factor is almost the same for both cases.

To reduce the fuel consumption rate of the pulse detonation combustor (PDC), an experimental study was carried out to study the influence of the injector aperture on the atomization particle size of the pneumatic atomization injector and its combustion characteristics of the PDC. The test used gasoline\air as the fuel and oxidant, tested the atomization particle size and obtained the detonation characteristics of PDC. The research results show that the atomization particle size of the injector under the three apertures decreases with the increase of the mixed air volume and increases with the increase of the fuel flow rate. The injector atomization effect is better when the aperture is 1.4 mm, the atomization particle size range is 71.8–119.8 μm. PDC using different injectors can achieve stable operation in the range of 5–20 Hz. Under the same working conditions, the smaller the injector aperture, the higher the average detonation pressure of the PDC.

In order to realize high-quality control for turboshaft engine with variable rotor speed, a cascade control method based on the acceleration estimator of gas turbine speed ( N gdot ) through Complementary filtering – incremental nonlinear dynamic inversion (CF-INDI)with Newton’s prediction is proposed. Firstly, an adaptive torsional suppression method based on the Recursive Least Square (RLS) algorithm and combined with the control index of power turbine speed is proposed and designed to solve the torsional instability of control system in turboshaft engine with variable rotor speed. Then, an online N gdot estimated method based on CF-INDI with Newton’s prediction is developed, and the complementary filtering method is applied to compensate for the inaccuracy of the state variable model available at a single working point, and Newton prediction is used to alleviate the problem of lagging acceleration estimate, which is accessible to realize the cascade control for turboshaft engine based on N gdot . The simulation results show that the RLS adaptive torsional filter can effectively suppress the low-order torsional components with frequencies of 1.2 Hz and 2.0 Hz mixing in power turbine speed, and the torsional suppression effect is remarkable. In addition, under different flight conditions, CF-INDI with Newton prediction is faster and more accurate than INDI N gdot estimate, and smoother than central difference method, which proves the feasibility of INDI N gdot estimator. Meanwhile, after the application of CF-INDI with Newton prediction N gdot estimator in cascade control, the anti-interference capacity of power turbine speed is stronger, and the dynamic control effect is superior, which has satisfactory robustness performance.

This study proposes two crescent cooling holes based on favorable vortex structures to enhance the film cooling effectiveness. Large eddy simulations and laser-induced fluorescence were conducted. An up-convex crescent hole (UCC) was established first, which uses an anti-counter-rotating vortex pair (anti-CRVP) to prevent the coolant from lifting off the surface and another CRVP to achieve better cooling uniformity at the centerline. Then, a down-convex crescent hole (DCC) was further optimized based on the vortex of the UCC. The results show that the film cooling effectiveness of the DCC increases by 48 % at BR = 1.6 compared to the UCC. Apart from the time-averaged results, the instantaneous flow field of DCC was analyzed in detail, and the distribution and breakdown of hairpin vortexes highly affect the cooling effectiveness. Besides, the heat transfer coefficient and net heat flux reduction were analyzed to overall evaluate the film cooling performance of DCC.

Mixing of Mach 2.1 circular, jet issuing from a straight convergent-divergent circular nozzle, in the presence of sonic air tabs at exit and shifted locations along the jet axis was investigated experimentally at nozzle pressure ratios (NPR) 3–6, insteps of 1. Two constant area tubes of 1 mm diameter positioned diametrically opposite, at 0 D, 0.25 D, 0.5 D and 0.75 D (where D is the nozzle exit diameter), were used for fluidic injection. The injection pressure ratio (IPR) of air tabs was maintained at 6. The Mach 2.1 jet operated at nozzle pressure ratio (NPR) in the range of overexpanded states corresponding to NPR 3–6 was controlled with the sonic air tabs operating at the underexpanded state corresponding to IPR 6. The impact of air tabs on jet mixing was studied from the measured Pitot pressure along the jet centerline. The centerline pressure decay of the jet confirms that the air tab promotes jet mixing with the entrained air mass, and the mixing promotion caused by the air tab is dependent on tab location as well as the NPR. In the presence of air tabs, the jet possesses shorter core and experiences faster decay than the uncontrolled jet. Also, the air tabs were effective in reducing the number of shock cells and rendering the waves weaker in the jet core. Among the tab locations, the mixing promoting effectiveness of air tabs at 0 D is better than the tabs at shifted locations. The jet core length reduction caused by the air tab at 0 D increases from 25.4 % to 77.2 %, with increasing NPR from 3 to 6. The same trend was noticed for tab location 0.75 D, but not for 0.25 D and 0.5 D locations. The core length reduction for 0.75 D tab location is about 61.4 %, at NPR 6, and 62.1 % and 55.8 %, for NPR 5 and 4 for tab locations 0.25 D and 0.5 D, respectively. Shadowgraph images of the waves present in the jet core confirms the findings of centerline pressure decay results.

A Subsonic Cruise Vehicle (SCV) is designed with a Submerged Air intake with appreciable total pressure recovery along with acceptable distortion levels. A steady, 3D, density based, one equation turbulence model in congregation with Navier Stokes is used to simulate the flow field in commercial CFD code. In the initial iteration of the intake, the numerical simulation pointed towards a large scope in the improvement of the geometry. Subsequently, the design changes have been successfully incorporated and analyzed. The performance of these modified intakes was evaluated by measuring distortion coefficient and total pressure recovery. An experimental validation case has been demonstrated in a subsonic wind tunnel, which proves the close match with the CFD prediction. The comparison between the performances of the intakes has been made. The redesigned submerged air intake is flown successfully with SCV at all flight conditions within the acceptable distortion limits.

An optimization methodology for a TBCC propulsion schemes was established for hypersonic vehicles, focusing on the integration of aircraft and engine performance. Altitude-velocity characteristics of TBCC propulsion were obtained through engine performance calculations. By analyzing mission requirements, lift-drag characteristics, and flight constraints, the take-off thrust-weight ratio and wing load of the vehicle were optimized to meet the flight conditions. In addition, the fuel ratio was calculated. To determine the vehicle’s gross take-off weight and the engine’s take-off thrust, a model considering the weight of the vehicle and the turbine engine were used. The optimization process selects four thermodynamic cycle parameters for the turbine engine as the independent variables. An improved particle swarm optimization-back propagation neural network was used to establish the relationship between the four parameters and the gross take-off weight, aiming to minimize the vehicle’s weight. The results of the optimization process show that the total take-off weight of the optimized vehicle has decreased from 112363.41 kg to 102218.98 kg. The required uninstalled take-off thrust of TBCC has also been reduced from 162.86 kN to 147.49 kN, resulting in a decrease in mass flow from 156.50 kg/s to 133.30 kg/s.

The current study aims to amplify the predictive ability of the numerical model developed for a gas turbine engine-based power plants by process of regeneration and intercooling. Artificial neural networks (ANN) and adaptive neuro-fuzzy interface systems (ANFIS) are the two techniques mainly concentrated in this study which were not properly implemented previously. The performance parameters namely, specific power (SP), thermal efficiency ( η ), and enthalpy based specific fuel consumption (EBSFC) of a Turboprop engine were predicted using thermodynamic parameters namely, pressure ratio (PR), nozzle pressure ratio (NPR), turbine inlet temperature (TIT), for constant regeneration (R), and intercooling (E) efficiencies. The results showed that a high regression result R 2 of 0.9831 and 0.9899 was found for the ANFIS model for η for training and testing, respectively. Also, the ANFIS model resulted in best performance of the performance characteristics when compared to ANN.

This paper numerically investigated the influence of bleed pressure on the shock-wave/boundary-layer interaction in a transonic compressor stator with a row of suction holes. These suction holes are arranged on the suction surface in the spanwise direction. Six different bleed pressure ratios are studied at an inlet Mach number of 1.0. Results show that the reduction in the pressure loss increases as the bleed pressure decreases, and the maximum relative decline is 36.4 % when the flow loss caused by the suction flow path is taken into account. The bleeding of suction holes changes the shock wave structure, and the modified shock wave structure is three-dimensional. As the bleed pressure decreases, the tangential momentum of boundary layer increases. A strong spanwise variation in the boundary layer is observed downstream of the suction hole. Turning the flow direction and creating a favorable pressure gradient are key reasons for controlling boundary layer separation.

Asymmetric leading edge offers several advantages over conventional leading edge. However, there are few related published research on compressor blades. This study proposes a blade parameterization method that superimposing the thickness distribution on the camber line at the middle and curve fitting at the leading and trailing edges. This parameterization method is applied to a stator in the final stage of an axial compressor, and numerical simulation is used to calculate the aerodynamic performance at various attack angles under the designed inlet Mach number. There is an optimal deflection angle regardless of which side of the leading edge deflects, and deflection toward the pressure side is conducive to the suppression of separation bubble near the leading edge at positive attack angle. When the maximum curvature of the pressure side is decreased, the operating range changed slightly, but the positive attack angle range grew from 5.166° to 5.466°.