Volume 36, Issue 2 -

New U.S.-DOD guiding instructions for 2019 JEST-R&D funded proposals, are now available, [1]. Past U.S.-decision “not to weaponize” JEST, has induced it to advance Anti-Gliding-Hypersonic-Scramjet-Threats, (GHST), [1, 2, 3, 4, 5, 6, 7, 8, 9, 10, 11, 12, 13, 14, 15, 16, 17, 18, 19, 20]. This editorial updates previous ones, [21, 22, 23, 24, 25, 26, 27, 28, 29, 30, 31], mainly on RE-EDUCATING JET-ENGINE RESEARCHERS TO STAY RELEVANT, especially in JETONAUTIC-POST-STALL-FLIGHT, where AOOF fails, [32, 33, 34, 35, 36, 37, 38, 39, 40, 41, 42, 43, 44, 45, 46, 47, 48, 49, 50, 51, 52, 53, 54, 55, 56, 57, 58, 59, 60, 61, 62, 63, 64, 65, 66, 67, 68, 69, 70, 71, 72, 73, 74, 75, 76, 77, 78, 79, 80, 81, 82, 83, 84]. Since 11.06.2018, China has joined the US-Russian jet-engines-steering (JES) global-race by JES-F-22 v. JES-Su-35 and 57 & JES-10B, [73, 74] v. THAAD, Arrow-3, David-Sling global network v. 6th-generation S500, Avangard and other systems, [85, 86].

The two-dimensional streamline curvature through-flow modeling of turbomachinery is still a key element for turbomachinery preliminary analysis. Basically, axisymmetric swirling flow field is solved numerically. The effects of blades are imposed as sources of swirl, work input/output and entropy generation. Although the topic is studied vastly in the literature for compressors and turbines, combined modeling of the transonic fan and the downstream splitter of turbofan engine configuration, to the authors’ best knowledge, is limited. In a prior study, the authors presented a new method for bypass fan modeling for inverse design calculations. Moreover, new set of practical empirical correlations are calibrated and validated. This paper is an extension of this study to rapid off-design analysis of transonic by-pass fan systems. The methodology is validated by two test cases: NASA 2-stage fan and GE-NASA bypass fan case. The proposed methodology is a simple extension for streamline curvature method and can be applied to existing compressor methodologies with minimum numerical effort.

To predict the pollutant emissions and lean blowout, chemical reactor network (CRN) model is applied to the modern aircraft engine combustion chamber. In this study, the CRN which represent the major features of aerodynamics and combustion in the combustion chamber is set up based on the OpenFOAM simulation results. The boundary and the initial conditions used for the CRN derive from the operating modes of typical aircraft engine cycle. A 21 species 30 steps chemical mechanism of kerosene is employed in the CRN method. The levels of pollutant emissions are obtained under four ICAO engine power settings of idle, approach climb and take off. The lean blowout equivalent ratio is evaluated at the idle power setting. The results will be helpful to predict the aircraft engine exhaust emissions and lean blowout (LBO).

We address requirements for laboratory testing of AC Dielectric Barrier Discharge (AC-DBD) plasma actuators for active flow control in aviation gas turbine engines. The actuator performance depends on the gas discharge properties, which, in turn, depend on the pressure and temperature. It is technically challenging to simultaneously set test-chamber pressure and temperature to the flight conditions. We propose that the AC-DBD actuator performance depends mainly on the gas density, when considering ambient conditions effects. This enables greatly simplified testing at room temperature with only chamber pressure needing to be set to match the density at flight conditions. For turbine engines, we first constructed generic models of four engine thrust-classes; 300-, 150-, 50-passenger, and military fighter, and then calculated the densities along the engine at sea-level takeoff and altitude cruise conditions. The range of chamber pressures that covers all potential applications was found to be from 3 to 1256 kPa (0.03 to 12.4 atm), depending on engine-class, flight altitude, and actuator placement in the engine. The engine models are non-proprietary and can be used as reference data for evaluation requirements of other actuator types and for other purposes. We also provided examples for air vehicles applications up to 19,812 m (65,000 ft).

The proper orthogonal decomposition (POD) method is employed to analyze the unsteady flow control mechanism because it is a good approach to decouple the spatial and temporal structures of unsteady flow fields. The results showed that the main effect of the periodic excitation is reallocating the energy of each mode, and selectively strengthening or weakening certain modes. Under proper amplitude and frequency of periodic excitation, the energy in higher modes will be transferred to the first mode and the translation of the modal energy is coming from the reconstructing of spatial flow structures and the ordering of modal evolution characteristics. The best control effect will be achieved when the total energy ratio of the first mode is the highest and the excitation frequency reaches the separation vortex frequency at the same time. In order to quantitatively analyze the order degree of the unsteady flow field, the maximum Lyapunov exponent was introduced. The results showed that with the energy in higher modes transferred to the lower modes, the flow field transfers from a disordered pattern to an ordered one.

In this paper the application of model predictive control (MPC) to a two-mode model of the dynamics of the combustion process is considered. It is shown that the MPC by itself does not stabilize the combustor and the control gains obtained by applying the MPC algorithms need to be optimized further to ensure that the phase difference between the two modes is also stable. The results of applying the algorithm are compared with the open loop model amplitude responses and to the closed loop responses obtained by the application of a direct adaptive control algorithm. It is shown that the MPC coupled with the cost parameter optimisation proposed in the paper, always guarantees the closed loop stability, a feature that may not always be possible with an adaptive implementations.

The quasi-one-dimensional method for the dual-mode scramjet (DMR) of the hypersonic RBCC powered vehicle was simplified in most of open researches. Furthermore, these simplified method can not fully capture the processes of wall heat transfer, changes in the boundary layer and the ratio of specific heat and the transonic flow in the reacting flow. Addressing this problem, we establish the models for processing core flow area, transonic flow and pre-combustion shock train (PCST) based on the governing equations for quasi-one-dimensional flow and certain assumptions. Thus the quasi-one-dimensional model of dual-mode scramjet engine that incorporates the changes in wall heat transfer and in the ratio of specific heat is built. Then, the reliability and accuracy of the model are assessed qualitatively and quantitatively by experiment and CFD numerical simulation. There is a high agreement between the theoretical calculations and the results of experimental data and CFD numerical simulation. This work expands the application scope and increases the reliability of quasi-one-dimensional model of dual-mode scramjet engine in RBCC engine. The results shed new light on the preliminary performance assessment and engineering application of dual-mode scramjet engine.

The forward bending centrifugal fan features long vanes and narrow passages, and the internal flow is apt to be affected by wall boundary layer of the blade surfaces. An aerodynamic investigation on the forward bending centrifugal fan is presented in this paper. The performances of the unsteady flow field in the centrifugal fan is predicted by solving the incompressible Reynolds-averaged Navier-Stokes (RANS) equations based on the Shear Stress Transport (SST) – Scale Adaptive Simulation (SAS) turbulence model. And three-dimensional (3-D) flow phenomenon of non-uniform fluid flow is discussed through observing the flow separation and vortex distribution in the impeller region. The installation location of inlet nozzle is finally adjusted based on the calculated results, and the centrifugal fan with eccentric inlet nozzle is proposed. From the results, the eccentric inlet nozzle with L=3 mm, θ=120° could improve the internal flow of the forward bending centrifugal fan and enhance its aerodynamic performance.