Band 34, Heft 1 -

The vortex valve solid variable thrust motor is a new solid motor which can achieve Vehicle system trajectory optimization and motor energy management. Numerical calculation was performed to investigate the influence of vortex chamber diameter, vortex chamber shape, and vortex chamber height of the vortex valve solid variable thrust motor on modulation performance. The test results verified that the calculation results are consistent with laboratory results with a maximum error of 9.5%. The research drew the following major conclusions: the optimal modulation performance was achieved in a cylindrical vortex chamber, increasing the vortex chamber diameter improved the modulation performance of the vortex valve solid variable thrust motor, optimal modulation performance could be achieved when the height of the vortex chamber is half of the vortex chamber outlet diameter, and the hot gas control flow could result in an enhancement of modulation performance. The results can provide the basis for establishing the design method of the vortex valve solid variable thrust motor.

Numerical analysis of the supersonic combustion and flow structure through a scramjet engine at Mach 7 with alternating wedge fuel injection and with three angle of attack ( α =−3°, α =0°, α =3°) have been studied in the present research article. The configuration used here is slight modification of the Rabadan et al. scramjet model. Steady two dimensional (2D) Reynolds-averaged Navier-Stokes (RANS) simulation and Shear stress transport (SST) based on k –ω turbulent model is used to predict the shock structure and combustion phenomenon inside the scramjet combustor. All the simulations are done by using Ansys 14-Fluent code. The combustion model used here is the combination of eddy dissipation and finite rate chemistry models since this model avoids Arrhenius calculations in which reaction rates are controlled by turbulence. Present results show that the geometry with negative angle of attack ( α =−3°) have lowest ignition delay and it improves the performance of scramjet combustor as compared to geometry with α =0°, α =3°. The combustion phenomena and efficiency is also found to be stronger and highest in case of α =−3°.

Failure of a high pressure (HP) turbine blade in a military turbofan engine is investigated to determine the root cause of failure. Forensic and metallurgical investigations are carried out on the affected blades. The loss of coating and the presence of heavily oxidized intergranular fracture features including substrate material aging and airfoil curling in the trailing edge of a representative blade indicate that the coating is not providing adequate oxidation protection and the blade material substrate is not suitable for the application at hand. Coating spallation followed by substrate oxidation and aging leading to intergranular cracking and localized trailing edge curling is the root cause of the blade failure. The remaining portion of the blade fracture surface showed ductile overload features in the final failure. The damage observed in downstream components is due to secondary effects.

The nozzle admittance is very important in the theoretical analysis of nozzle damping in combustion instability. The linearized Euler equations (LEE) are used to determine the nozzle admittance with consideration of the mean flow properties. The acoustic energy flux through the nozzle is calculated to evaluate the nozzle damping upon longitudinal oscillation modes. Then the parametric study, involving the nozzle convergent geometry, convergent half angle and nozzle size, is carried out. It is shown that the imaginary part of the nozzle admittance plays a non-negligible role in the determination of the nozzle damping. Under the conditions considered in this work ( f* =1,000 Hz, de *=0.18 m), the acoustic energy flux released from the nozzle with a 30 o convergent half angle is highest (30 o :6.0×104kgs−3$6.0 \times {10^4}{\rm{kg}}{{\rm{s}}^{- 3}}$, 45 o :5.2×104kgs−3$5.2 \times {10^4}{\rm{kg}}{{\rm{s}}^{- 3}}$, 60 o : 4.9×104kgs−3$4.9 \times {10^4}{\rm{kg}}{{\rm{s}}^{- 3}}$). The change of nozzle convergent geometry is more sensitive for the large size nozzle to increase the nozzle damping.

The turbofan engine operation consists of a number of hardware and software. The engine is controlled by Electronic Engine Control Unit (EECU). In order to control the engine, EECU communicates with an aircraft system, Actuator Drive Unit (ADU), Engine Power Unit (EPU) and sensors on the engine. This paper tried to investigate the process form starting to taking-off and aims to design the EECU software mode and defined communication data format. The software is implemented according to the designed software mode.

The aim of the present paper is to investigate the propulsive performance of the hollow rotating detonation engine (RDE) with a Laval nozzle. Three-dimensional simulations are carried out with a one-step Arrhenius chemistry model. The Laval nozzle is found to improve the propulsive performance of hollow RDE in all respects. The thrust and fuel-based specific impulse are increased up to 12.60 kN and 7484.40 s, respectively, from 6.46 kN and 6720.48 s. Meanwhile, the total mass flow rate increases from 3.63 kg/s to 6.68 kg/s. Overall, the Laval nozzle significantly improves the propulsive performance of the hollow RDE and makes it a promising model among detonation engines.

A case of compressor surge with bang noise during takeoff roll is investigated and presented in this paper. Fatigue failure of compressor rotor blades during takeoff is found to disturb the aerodynamics of compressor flow causing the surge. Based on evidences, failure of rotor blades and compressor surge due to over-speed and foreign object debris is ruled out. The paper presents the methodology adopted for the investigation and also suggests remedial measures necessary to prevent such incidents.

The combined effects of upstream wakes and surface roughness on boundary layer development have been investigated experimentally to improve the performance of ultra-high-lift low-pressure turbine (LPT) blades. The measurement was performed on a linear cascade with an ultra-high-lift LP turbine profile named IET-LPTA with a Zweifel loading coefficient of about 1.4. The wakes were simulated by the moving cylindrical bars upstream of the cascade. The surface roughness was achieved using sandpaper strips which were placed into the slot incised on the blades surfaces. Three types of slots combined with three types of roughness heights formed a large measurement matrix. The roughness with a height of 8.82 μm (1.05×10 −4 chord length) covering 5.2 % suction surface reduced the profile loss in the whole tested Reynolds number range. The details of the boundary layer were measured with a boundary hot-wire probe. It was found that both the wake-induced transition and the nature transition between wakes were hastened due to the K-H instability expanded by the roughness of the S3-R3 surface. The nature transition process was more sensitive to surface roughness than the wake-induced transition.

Non-reactive flows moving through a model scramjet were investigated using a delayed detached eddy simulation (DDES), which is a hybrid scheme combining Reynolds averaged Navier-Stokes scheme and a large eddy simulation. The three dimensional Navier-Stokes equations were solved numerically on a structural grid using finite volume methods. An in-house was developed. This code used a monotonic upstream-centered scheme for conservation laws (MUSCL) with an advection upstream splitting method by pressure weight function (AUSMPW+) for space. In addition, a 4th order Runge-Kutta scheme was used with preconditioning for time integration. The geometries and boundary conditions of a scramjet combustor operated by DLR, a German aerospace center, were considered. The profiles of the lower wall pressure and axial velocity obtained from a time-averaged solution were compared with experimental results. Also, the mixing efficiency and total pressure recovery factor were provided in order to inspect the performance of the combustor.

In order to realize the stator’s internal flow field measurement of the multi-stage axial compressor, many different lengths of L-type five-hole probes and one four-hole probe have been designed. The detailed 3D (three-dimensional) flow fields in the stator blade passage of original and modified large scale compressors have been measured with the probes traversed by a probe traverse mechanism. The objective of the study is to assess the measurement method of adopting multi-hole pneumatic probes to achieve the flow field of stator vane passage. Results clearly demonstrate the flow field characteristics of stator blade passage in original and modified compressors. Furthermore, the reliability of adopted multi-hole pneumatic probes is validated. Therefore, it is proved that the probe traverse mechanism drives the L-type five-hole probe or other probe method can be applied to internal flow field measurements for the stator of large scale multi-stage compressors.

The combustion at high speed reactants requires a flame holding characteristics to sustain the flame in the afterburner. The flame holding characteristics of the combustor is carried out by the bluff-body stabilizers. The range of conditions of parameters influencing the flame stabilization is to be identified and the effects on the flame sustainability have to be investigated. DeZubay used the concept of DeZubay number and flame stability envelope to determine the stabilization and blowout range. In the present work, the effect of air pressure and the angle of apex of the V-gutter on flame stabilization and blowout mechanism have been experimentally investigated for six different apex angles and four different air pressure conditions. The value of DeZubay number at each condition has been calculated and verified with DeZubay stability chart for flame stabilization. The results show that stable flame is obtained for the entire pressure range when the apex angle of the V gutter is in 60° and 90°.

Aerodynamic mixing of subsonic and sonic jets with limiting tabs, with and without corrugations, has been studied experimentally. Limiting tab located at the nozzle exit and at a downstream distance of 0.5D has been considered in this study. Mixing caused by the tab at nozzle exit is found to be better that of tab at 0.5D, for both plain and corrugated geometries. Also, both tabs caused better mixing for underexpanded sonic jets than the correctly expanded sonic jet and subsonic jets. At nozzle pressure ratio 3 the plain tab at the nozzle exit reduced the core by about 56 % and the corrugated tab by about 51 %. But when the plain tab is placed at 0.5D the jet mixing is retarded. However, the corrugated tab at 0.5D enhances the mixing, though not up to the level of the same tab at 0D, at all Mach numbers except 0.6. The maximum reduction of core caused by shifted corrugated tab is 14 % for Mach 0.8 jet.