Volume 35, Issue 4 -

I often open my lectures with a question: Do you agree that as speed goes down in any landing , the vertical tail fails to orient the nose back to the runway under strong side winds while ONLY side deflecting jet-engine-steering ( JES ) can safely re-orient it to safe landing? Just facts, less opinions; And if you do, you accept the 1ST-element of jetonautics defined by 12 figures below, text and references 1 to 48.

The forthcoming 2020-2040 6 TH -Generation is characterized by a host of debated and confused CASLAP-strategies, doctrines, tactics, mission definitions, jetonautic-funding, systems-of-systems technologies vis-à-vis AOOF-based-prop-swarm-technologies and almost symmetric jetonautics at both sides, [Part-A references 5,10,11,17,20,24,26,32,36,48,49] . CASLAP offensive-defensive options for each country vary with jetonautics, location, time, scientific and technological education, economics, geopolitics and stealth/super-agility/armament/AI of its fighter aircraft, UCAVs, JES-UCAV, prop-drones or USV, LCS, ULV, each with a different control-communication and various propulsion systems ranging from $100/unit in prop-drone-swarms to $10+ million per jet engine unit with spares, parts, maintenance, etc. It therefore deems timely to introduce an initial degree of order and classification to prevent today future misunderstanding/errors in designs, funding and encounters – the central aim and scope of these editorials.

In this work, a hydrogen fueled dual-mode scramjet combustor was investigated experimentally. Clean and dry air was supplied to the combustor through a Mach 2 Nozzle with a total temperature of 800 K and a total pressure of 800 kPa. The high enthalpy air was provided by an electricity resistance heater. Room temperature hydrogen was injected with sonic speed from injector orifices vertically, and downstream the injector a tandem cavity flame holder was mounted. Except wall pressure profiles, velocity and temperature profiles in and at exit of the combustor were also measured using hydroxyl tagging velocimetry (HTV) and tunable diode laser absorption spectroscopy (TDLAS) respectively. Results showed that combustion occurred mainly at the bottom side of the combustor. And there were also an extreme disparity of the velocity and temperature profiles along the Y direction, i.e. the transverse direction.

Isolator is a critical component in supersonic air breathing engine and it is usually situated between the inlet and the combustor of a dual-mode ramjet/scramjet engine. In the present study, shape transition effects on isolator performance have been studied by carrying out experimental investigations on square, square to circular and square to elliptical transition ducts. The length of the isolator chosen in this study is 180 mm and the cross-sectional area of 900 mm 2 is maintained constant along the length for all the ducts. Experiments were carried out for isolator inlet Mach 2, using a contoured nozzle. Varying the pressure of the settling chamber varied the expansion level at the nozzle exit, which run the nozzle. The wall static pressure along the length of the isolator and the Pitot pressure at the exit plane of the isolator were measured for all the configurations. Shadowgraph technique was employed for visualizing the shock-train in the isolator. The square to circular transition isolator is found to be more efficient in achieving the static pressure rise across the isolator than the square and square to elliptical transition ducts.

The effects of different inlet parameters such as inlet temperature and pressure on combustion performance in a single-head combustor were experimentally investigated in this study. The combustion efficiency, total pressure loss, and CO and NO emissions at the outlet of a single-head rectangular combustor with different types of swirlers were separately measured. The experimental results showed that the inlet parameters had obvious effects on the combustion performance, with critical values of 600 K for the inlet temperature and 3.5 bar for the inlet pressure. The combustion efficiency noticeably increased with an increase in the inlet pressure or temperature below these values; however, when either of the inlet parameters was above the critical value, the combustion efficiency was approximately 100 %; that is, the combustion efficiency changed little with an increase in inlet temperate or pressure. When the inlet temperature or pressure increased, NO emission increased but CO emission decreased. By fitting curves to analyze the experimental data, the empirical relationships between the emissions and the inlet temperature were observed to be CO∝e−T,NO∝eT$CO\; \propto \;{e^{ - T}}, NO\; \propto \;{e^T}$, and those between the emissions and the inlet pressure were CO∝ea+bP+cP2,NO∝eP$CO\, \propto \,{e^{a + bP + c{P^2}}}, NO\, \propto \,{e^P}$. The total pressure loss increased with the inlet temperature.

Various model-based methods are widely used to aircraft engine fault diagnosis, and an accurate engine model is used in these approaches. However, it is difficult to obtain general engine model with high accuracy due to engine individual difference, lifecycle performance deterioration and modeling uncertainty. Recently, data-driven diagnostic approaches for aircraft engine become more popular with the development of machine learning technologies. While these data-driven methods to engine fault diagnosis tend to ignore experimental data sparse and uncertainty, which results in hardly achieve fast fault diagnosis for multiple patterns. This paper presents a novel data-driven diagnostic approach using Sparse Bayesian Extreme Learning Machine (SBELM) for engine fault diagnosis. This methodology addresses fast fault diagnosis without relying on engine model. To enhance the reliability of fast fault diagnosis and enlarge the detectable fault number, a SBELM-based multi-output classifier framework is designed. The reduced sparse topology of ELM is presented and utilized to fault diagnosis extended from single classifier to multi-output classifier. The effects of noise and measurement uncertainty are taken into consideration. Simulation results show the SBELM-based multi-output classifier for engine fault diagnosis is superior to the existing data-driven ones with regards to accuracy and computational efforts.

The effects of turbulent burning velocities in a turbulent premixed combustion simulation with a G-equation are investigated using the 3D LES technique. Two turbulent burning velocity models – Kobayashi model, which takes into account the burning velocity pressure effect, and the Pitsch model, which considers the flame regions on the premixed flame structure – are implemented. An LM6000 combustor is employed to validate the turbulent premixed combustion model. The results show that the flame structures in front of the injector have different shapes in each model because of the different turbulent burning velocities. These different flame structures induce changes in the entire combustor flow field, including in the recirculation zone. The dynamic mode decomposition (DMD) method and linear acoustic analysis provide the dominant acoustic mode.

Modern day gas turbines are prime movers in land, air and sea. They have stringent performance requirements to meet the complex mission objectives. Optimal control strategies can help them meet their performance objectives more efficiently. A novel inverse simulation method for optimal control and system analysis studies using Differential Algebraic Equality/Inequality (DAE/DAI) technique is brought out in this paper with a case study. The gas turbine model together with safety constraints and performance specifications is represented as a high index DAI/DAE system. The solution for this DAE/DAI system is obtained using a new numerical approach that is capable of handling both equality and inequality constraints on system dynamics. The algorithm involves direct numerical integration of a DAI formulation in a time stepping manner using Sequential Quadratic Programming (SQP) solver that detects and satisfy active constraints at each time step (mesh point). In this unique approach the model and the constraints are always solved together. The method ensures stable solution at each time step, local minimum at each iteration of simulation and provides a regularised basis to the solver. Compared to other existing computationally intensive techniques in usage, this approach is easy, ensures continuous constraint satisfaction and provides a viable option for Model Predictive Control (MPC) of gas turbine engines.

This paper aims at optimizing a Turbine Based Combined Cycle (TBCC) nozzle for upgrading its aerodynamic performance in multiple flight conditions. An in-house RANS solver called NSAWET is employed for aerodynamic evaluation. The optimizer is a differential evolution algorithm combined with a response surface. Firstly, a two-dimensional model of the initial TBCC nozzle system is investigated. The flow field of the nozzle contains complicated shockwave interactions that cause thrust loss. Then multi-point aerodynamic optimization of a two-dimensional ramjet nozzle is carried out, which objectives are to maximize the thrust coefficients at Mach numbers 2.5, 3.0 and 4.0. The objective functions are increased substantially at the first two Mach numbers after optimization, while slightly decreased at the last Mach number. The turbine flow path is built up based on the optimized profile and the performances in typical flight conditions are validated. Results demonstrate that both the ramjet and turbine nozzles are improved in most flight conditions.

The main objective of the present paper is to provide an approximate method for nonlinear system with a family of equilibrium points (EPs). By investigating the system’s equilibrium manifold (EM) and its expansion form, a class of nonlinear system is modeled analytically. The property of the EM expansion form and the effect of mapping design and parameterizing method to the model have been discussed. Then an approximate nonlinear model for aircraft engine was applied, followed by an identification procedure for aircraft engine. It is shown that the modeling method based on system’s EM can not only comply with the system physical characteristics, but also satisfy the accuracy of modeling requirement. Numerical simulations are given to demonstrate the good precision in capturing the nonlinear behavior of nonlinearities and a simpler structure.

The optimization of turbine cooling design has become a new research field of gas turbine. The swirl chamber is a prospect cooling concept. In this paper, the numerical simulation of the swirl chamber is carried out by FLUENT. The influence of inlet size parameters, temperature ratio and inlet Reynolds number on the enhanced heat transfer of swirl chamber is studied. The results show that, in the range of the studied condition, Nusselt number decreases with the height, the width, the ratio of width to height and Reynolds number. It also shows that comprehensive heat transfer effect is best at d =20 mm and enhances observably with the enlargement of width, width height ratio, and Reynolds number. Friction factor increases with height, width, temperature ratio and Reynolds number decreases. It is increased by increasing width height ratio. Nusselt number and comprehensive heat transfer effect decrease a little with aggrandizement of temperature ratio.