Band 42, Heft 1

A computational study of a three ramped dual duct Rocket Based Combined Cycle (RBCC) engine inlet at scramjet mode using different types of MVG (micro vortex generator) arrays were conducted. The definite geometry of engine inlet was operated at hypersonic speeds of Mach 5 and 7 to study the effect of the arrays of delta ramp (DR), rectangular vane (RRV) and ramp vane (RV) on the pressure recovery, exit Mach number, the mass flow rate and Shock wave Boundary Layer interaction (SWBLI). The study was performed considering same heights for all the configurations of MVG array and were positioned at the point of shock impingement on the ramp which caused the separation bubble. The computational analysis was done using k-omega model in Fluent Workbench of Ansys.

A major focus of the present study is to construct high-fidelity models for predicting corrected mass flow rates based on the collected compressor map data. Both traditional machine learning research and modern deep learning techniques have been employed to obtain well-fitted regression models of compressor mass flow rate. As traditional machine learning methods, Multiple Linear Regression and Random Forest, are conducted on compressor maps for prediction of corrected mass flow rate. The time series-based deep learning models are able to capture the overall trend of a given input for specific map data. Therefore, a time series-based deep learning technique, namely Gated Recurrent Unit has been employed to improve regression results. Besides, the prediction capabilities of the models, results also show that the proposed models can be used for the development of dynamic aero-thermal mathematical models of gas turbine engines and mass flow rate models created for dynamic compressors in other disciplines.

Auxiliary systems, such as regeneration and intercooling, have been integrated with the primary gas generator to improve power production and fuel economy in modern gas turbine power plants. Implementing these techniques in turbine engines is challenging due to size, weight, and complex flow patterns. A solution is to use a turboprop engine with a smaller mass flow rate and simpler gas paths. The current study involves the numerical analysis of performance parameters namely, specific power (SP), thermal efficiency ( η ), and enthalpy based specific fuel consumption (EBSFC) of a turboprop engine using thermodynamic parameters namely, pressure ratio (PR), nozzle pressure ratio (NPR), turbine inlet temperature (TIT), regeneration efficiency (R), and intercooling efficiency (E). The results prove that the introduction of regeneration and intercooling showed significant improvement in the power developed, and reduced fuel consumption.

A Pulse Detonation Ramjet (PDR) performance evaluation model is established based on the component-based method. The influence law of working parameters of pulse detonation combustor (PDC), PDR geometry parameters, flight speed and altitude on the performance is analyzed. Results are as follows. (1) With the increase of the equivalent ratio, the PDR specific thrust first increases and then decreases, reaching a maximum value around 1.2, but the fuel specific impulse always decreases; (2) The change of inlet throat cross-sectional area has little effect on specific thrust and fuel specific impulse, but the increase of nozzle throat cross-sectional area will have a negative effect; (3) PDR specific thrust and fuel specific impulse increase first and then decrease with the increase of flight Mach number and altitude; (4) Compared with a ramjet engine, the advantage range of PDR can reach 18.04 %, but gradually decreases with the increase of Mach number.

Gas turbine engines are used extensively in a variety of industries, including power generation, aviation, and marine propulsion. One of the most difficult issues in developing gas turbine engines is dealing with the high temperatures. Internal cooling is a typical strategy for keeping important components, such as turbine blades, at a safe working temperature. Pin-fin cooling is a significant technology used in gas turbine blades to manage severe operating temperatures. This research paper investigates the impact of a new geometry combining two pin fins in cooling channels in gas turbine blades. The study involved testing seven different pin-fin arrays over a Reynolds number varying from 10,000 to 80,000 and a constant heat flux of 3280 W/m 2 applied to all the surfaces. The findings indicate that the use of pencil pin fins leads to a notable enhancement in cooling efficiency owing to the existence of a low-pressure drop, 5.2 % compared to pencil and tear drop pin fins and by 1.6 % compared to pencil and oblong pin fins, and it is 1.7 % less than pencil and circular fins.

The mixed compression scramjet engine intake was numerically simulated to study the characteristic of shock wave induced boundary layer separation leading to the formation of the separation bubble (SB). The analysis employed a 2D-RANS method with SST k-ω turbulence model at different Mach numbers. The intake is designed as a three-ramp intake for improved performance. As the Mach number varies, the size of the separation bubble formed due to the interaction of oblique shock waves with the boundary layer also varies, affecting both intake the efficiency and overall efficiency of the engine. Apart from traditional control techniques, the most preferred bleed technique is incorporated. Localised & distributed bleed techniques are designed and analysed within the intake at different locations. These techniques result in a reduction in the size of the separation bubble within the intake. Establishing perforation in the engine intake also increases the intake efficiency and overall engine performance.

The uneven temperature distribution at the combustion chamber outlet seriously affects the working life of the engine. In order to reduce the heat spot temperature at the combustion chamber outlet, a performance optimization control method of the engine minimum heat spot temperature pattern is proposed. Firstly, based on CFD method, the temperature distribution characteristics of combustion chamber outlet under different working conditions were obtained, and a component-level model of turbofan engine was established to characterize the heat spot temperature at combustion chamber outlet. Secondly, the high precision and high real-time engine on-board model is established by deep neural network. Compared with the component-level model, the average relative error of each performance parameter is less than 0.3 %, and the real-time performance is improved by 12 times. Finally, based on the feasible sequential quadratic programming algorithm, the performance optimization control of the minimum hot spot temperature model in the typical flight envelope is simulated and verified. The simulation results show that under the condition of ensuring the safe and stable operation of the engine and constant thrust, the heat spot temperature at the combustion chamber outlet decreases by 21 K maximum. Compared with the conventional minimum turbine front temperature optimization mode, the minimum heat spot temperature mode significantly reduces the heat spot temperature at the combustion chamber outlet.

To optimize the integrated flameholder, PIV was used to study flow fields of V-gutter and integrated flameholder under both non-reacting and reacting conditions. PLIF, high-speed cameras, and TDLAS were adopted to capture OH distribution, flame structure, and temperature distribution. Comparative analysis of flow fields, combustion characteristics and flame stabilization mechanisms were analyzed. Results show that heat release increases adverse pressure gradient, which can enlarge the recirculation zone size and recirculation rate compared to non-reacting flow field. The flames of both flameholders exhibit symmetrical structures distributed near the shear layers. The blockage ratio dominates the non-reacting flow field, while the expansion angle dominates the reacting flow field, which can further increase the adverse pressure gradient under reacting condition. The V-gutter flameholder demonstrates better fuel/air mixing and larger recirculation than the integrated flameholder. The combustion performance of the integrated flameholder is inferior to the V-gutter flameholder, albeit with better flow resistance properties.

The improvement of aero-engine performance has posed a serious threat to aeroelastic stability, thereby compromising the reliability of aero-engines. An effective approach to quantify the risk of compressor blade instability and enhance aeroelastic stability is through flutter probability evaluation. This study proposes a prediction method called the Particle Swarm Optimization-Deep Extremum Neural Network model (PSO-DENN) to improve the modeling accuracy and computational efficiency of compressor blade flutter probability analysis in aero-engines. Through deterministic analysis, the flutter response distribution of the blade is obtained. To account for the randomness of boundary conditions and time-varying loads, the flutter reliability of compressor blades is evaluated, providing insights into distribution characteristics, and reliability associated with aeroelastic instability. Comparative analysis of different methods demonstrates that the proposed PSO-DENN method improves calculation efficiency while ensuring accuracy.

Different types of geometric variations often appear coupledly in manufactured blades. It is desired to identify the ones that have the strongest impact on the performance. In this paper, the influence of multiple geometric variations on the compressor rotor blade at the design point is studied. Two hundred and nine varied blades are constructed by adding variation data to the design-intent blade and assessed using steady-state Reynolds-Averaged Navier–Stokes (RANS) simulations. It is shown that the region near the lower and upper end of the blade is more sensitive to geometric variations. Spearman’s rank correlation coefficient is used to measure the sensitivity of rotor performance to geometric variations. The results show that the stagger angle variation is the most influencing variation. Local sensitivity analysis at different spans also shows that the stagger angle variation is an important geometric variation that needs more attention during blade manufacturing. Also, the influence of geometric variations can reach further than nearby regions. Consequently, experiments on linear cascades at different spans of the rotor are carried out to study the effect of stagger angle on flow characteristics. Results show that the stagger angle variation could lead to different changes in performance depending on the specific cascade profile.

A novel NMPC (Nonlinear Model Predictive Control) based on composite predictive model is proposed and applied to direct engine thrust control. To improve the real-time of NMPC, an adaptive composite model based on SVM (State Variable Model), KF (Kalman Filter), and CLM (Component Level Model) is proposed as predictive model. The correction theory is adopted to establish a full envelope adaptive on-board predictive dynamic model and reduce the data storage of predictive model. At each sampling time, the CLM is calculated only once in the proposed NMPC, instead of many times in the popular NMPC based on EKF (extended Kalman filler). Therefore, the proposed NMPC has better real-time performance than the popular one. The simulations that consist of the proposed NMPC, the popular NMPC based on EKF, and the traditional controller PID are conducted. The simulations demonstrate that the proposed NMPC not only has greatly better real time performance than popular NMPC, but also has faster response speed than traditional controller PID.

Turbine-based combined cycle (TBCC) propulsion system is becoming one of the most promising propulsion systems for two-stage-to-orbit reusable launch vehicle. Mathematic model of this combined cycle engine is helpful for the basic understanding of performance analysis of this propulsion system. We developed mathematic model of TBCC propulsion system based on C++ platform in this paper. Firstly, turbojet engine was built on component level and ramjet engine was calculated through stream thrust function. Then, performance of turbine-based combined cycle propulsion system along a specific flight trajectory was investigated. According to the thrust of this combined cycle engine, mode transition point was suggested at Ma 2.5, which may achieve smooth mode transition from turbine mode to ramjet mode. Finally, mode transition based on smooth mass flow and smooth thrust criteria were studied. The thrust gap arises during smooth mass flow mode transition, particularly when the turbojet engine afterburner is powered off at the start of the mode transition, and it meets 34 % of the total thrust. Smooth thrust mode transition may be achieved by delaying the turbojet engine afterburner power off and injecting additional fuel into ramjet burner.

In this study, the flow field and performance for the aluminum based water ramjet was analyzed through numerical simulation. The time decrease in evaporation heat absorption process expands the range of aluminum/water combustion and increase the specific impulse. If the diameter of atomized water droplets from the primary inlet is large, the evaporation efficiency will be further reduced, which is not conducive to the further propagation of the reaction. However, if the diameter is small, the flow field temperature will be rapidly dropped due to rapid evaporation, and a large range of low-temperature areas will appear in the reaction area of the combustor. The diameter reduction of atomized water droplets from the secondary water inlet is conducive to the diffusion of the reaction area downstream, further and makes the combustion range be expanded. This effect is benefit for increasing the specific impulse and can reduce the temperature near the wall.

This study uses Multi-Island Genetic Algorithm (MIGA) and three-dimensional Computational Fluid Dynamics (CFD) software to optimize butterfly-shaped film cooling holes in the upper-stage rocket engine thrust chamber. The goal is to meet thermal protection and thrust requirements at high altitudes without re-ignition. To facilitate an all-encompassing worldwide search, the holes in the optimized design remain at set dimensions. Film continuity and stability at the nozzle outlet are greatly impacted by the hole structure. Inlet and divergence angles have little effect on thrust, according to regression research, but lip height ( de ) and outlet width ( β ) have a big impact on cold gas ejection, which affects cooling and thrust. Optimized results lead to a 20.49 K decrease in the monitoring section’s average wall temperature and a 52.8 N boost in thrust by reducing interference between supersonic airflow and extending film stability.

For the methanol jet atomization in a co-flow gas lifting flame, the self-developed WAVE model and KH-RT model were employed using the Saturne program to simulate the primary breakup of the jet column and the atomization of spray particles. Additionally, the unsteady flamelet model was utilized to simulate the combustion process. Comparing the calculated flame lifting height, droplet diameter distribution, and velocity distribution with experimental results demonstrates the atomization model’s ability to accurately capture the methanol fuel atomization process. This comparison verifies the reliability of both the atomization and combustion models, indicating their high accuracy. The study shows that upon injection from the nozzle, methanol fuel rapidly diffused, creating a cone-shaped spray structure. The high concentration of OH radicals is located close to the boundary between the regions of OH radical distribution and methanol distribution, gradually spreading downstream of the methanol jet. Furthermore, an increase in fuel flow rate resulted in a reduction in the flame’s lift height.

The study is aimed at obtaining and analyzing the flow and cavitation characteristics of the proposed control devices, which are in demand in control systems for aircraft and rocket engines, chemical and other technological processes that require a wide range of flow control at high pressure drops. The characteristics of vortex throttling devices with cylindrical and spherical chambers were obtained, providing flow control in the range of 0.1–25 kg/s. A change in flow rate in the range of 3.5–3.8 can be achieved by changing the flow swirl intensity without throttling the flow sections. When vortex devices operate, the threshold for the occurrence of cavitation shifts to the region of small values of P out /P in , and, starting from the values of the geometric characteristic A  > 4.5, cavitation is absent throughout the entire range of flow control. The pressure drop at the inlet channel of the vortex chamber does not exceed 10–30 % of the total resistance of the device. The torque on the control roller is small, 1.8–2.8 Nm, regardless of the flow rate.

Transonic axial compressors are widely used in aero-engines and gas turbines, and their efficiency has a significant influence on the equipment energy utilization rate and fuel cost. In this research, blade section modification based on flow analysis is carried out on the Rotor 67 for improving efficiency. First, the numerical calculation method used is verified with experimental data to ensure the accuracy of the simulation. Then, according to the analysis of simulated flow field and shock structure, the blade section shapes are altered to optimize the blade passage wide distribution for the sections from hub to mid-span. In addition, the blade surface angle is altered to reduce the supersonic flow expansion angle for decreasing shock strength. With these optimizations, the efficiency of modified transonic rotor increases by 0.9 percentage at the design point and 1.0 percentage at near stall point ( N  = 1.0), respectively. Slightly higher efficiency at off-design speeds ( N  = 0.9 and 0.8) are also obtained, with maintaining the stall margin.

This current work presents the performance analysis of a supersonic combustor interface and flow construction through a scramjet engine with a planar shaped strut injector (PSSI) at the supersonic Mach. An important aspect of this study is discovering the fuel mixing mechanism inside the combustion chamber with PSSI. The novel PSSI configuration can enhance mixing and combustion performance. The scramjet configuration is incorporated with a supersonic inlet air temperature of 1250 K, where the vitiated air follows at Mach 3, and this technique is based on a high accelerating effect of the scramjet propulsion mechanism. Scramjet engines can maintain naturalistic high enthalpy conditions in minimum durations. It is observed that the maximum temperature of 3510 K is attained at the recirculating zones produced because of undulation enlargement and therefore the fuel jet losses concentration whereas the maximum combustion efficiency of 86 % is investigated from the current research work.

Submerged air intake is often acquainted with the merit of reducing the drag and fitting well within the frame of weight constraints imposed on subsonic cruise vehicle. A novel method for designing an Intake was developed by GTRE, which has been successful with the validation against wind tunnel testing. The designed intake has been integrated with engine bay area of subsonic cruise vehicle. In this study, a provision of bleed from the ramp surface of the intake into the bay through splitter is used for cooling the engine frame, mechanical fixtures, oil tank and other electronic instruments mounted around the engine frame. Different sets of cowl ventilation slots provided on the bay wall will effuse this by-pass air from the intake into the atmosphere there by influencing the heat distribution within the bay area. The cascading effect of the intake with the bay area is extensively studied in this paper.