Search

This work studies the lift enhancement effect with a channel wing propeller configuration. Numerical simulations and flow field analysis are performed to investigate the influence (the relative position and spacing, angles of attack (AoA), propeller speed, incoming flow speed) on the lift enhancement effect. The aerodynamic performance is numerically simulated by the method of multi-reference frame (MRF). Numerical results demonstrate that the propeller suction process has a very significant lifting effect. The lifting effect can reach 5 times contrast to the single channel wing. The propeller located at 100% of the chord length of the channel wing achieves the best lift enhancement effect. The intervention of the propeller delays the flow separation at the trailing edge of the channel wing. It holds a high lift coefficient at a high AoA. The lift line slope remains essentially unchanged in the range of 0–30°. It is roughly the same as that of the single channel wing configuration before stall. With the increase of propeller speed or incoming flow speed, the lift coefficient increases significantly. The lift line slope is approximately the same at different speeds. The regularity functions of different coefficient are obtained through mechanism analysis and least squares data fitting methods. The lift coefficient changes with the AoA in a linear relationship and with the propeller speed or incoming flow speed in a quadratic relationship. [ABSTRACT FROM AUTHOR]

This work studies the lift enhancement effect with a channel wing propeller configuration. Numerical simulations and flow field analysis are performed to investigate the influence (the relative position and spacing, angles of attack (AoA), propeller speed, incoming flow speed) on the lift enhancement effect. The aerodynamic performance is numerically simulated by the method of multi-reference frame (MRF). Numerical results demonstrate that the propeller suction process has a very significant lifting effect. The lifting effect can reach 5 times contrast to the single channel wing. The propeller located at 100% of the chord length of the channel wing achieves the best lift enhancement effect. The intervention of the propeller delays the flow separation at the trailing edge of the channel wing. It holds a high lift coefficient at a high AoA. The lift line slope remains essentially unchanged in the range of 0–30°. It is roughly the same as that of the single channel wing configuration before stall. With the increase of propeller speed or incoming flow speed, the lift coefficient increases significantly. The lift line slope is approximately the same at different speeds. The regularity functions of different coefficient are obtained through mechanism analysis and least squares data fitting methods. The lift coefficient changes with the AoA in a linear relationship and with the propeller speed or incoming flow speed in a quadratic relationship. [ABSTRACT FROM AUTHOR]

Unsteady turbulent separated flow widely exists in aerospace vehicles and has become a major research topic. For the precise turbulent flow prediction, efficient and accurate numerical methods are the premise and foundation. However, most shock-capturing schemes designed for compressible flows have some difficulties in low-speed flow regions, and reduce the reliability when simulating unsteady turbulent separated flow with rich low Mach number features. Thus, the paper proposes a high-fidelity low-dissipation scheme termed LD-Roe2 and investigates the performances of the novel approach in unsteady turbulent separated flow simulations using the Hybrid Reynolds Averaged Navier-Stokes Equations/Large Eddy Simulation Approach (Spalart-Allmaras Delayed Detached Eddy Simulation). Numerical results including flow over a circular cylinder at subcritical Reynolds number and flow over tandem cylinders, demonstrate that the novel low-dissipation scheme can better simulate the unsteady separated flow and depicts finer smaller-scale vortical structures compared with traditional numerical schemes. It is expected to provide a more accurate and applicable numerical solution for the refined turbulent prediction. [ABSTRACT FROM AUTHOR]

In the early stage of rocket hot separation, the vented plume from separation gap interacts with the supersonic incoming flow, causing lateral interference which affects separation safety and flight stability. In this work, steady interstage interference flow fields of a two-stage rocket in different incoming flow conditions (Mach number from 2.5 to 5.5, angle of attack from 0° to 6°, flight altitude from 16 Km to 24 Km) are numerically studied by the compressible Reynolds Averaged Navier-Stokes equations with SST turbulence model. An annular lateral jet is used to simulate the vented plume near the separation gap. Firstly, the interference flow fields with and without angle of attack are analyzed in detail to illustrate the interference mechanism of lateral jet. Then, the effects of flight altitude, jet total pressure, separation distance and free-stream Mach number on lateral jet interaction are investigated. With the increase of flight altitude, jet total pressure and separation distance, the upstream separation length and the pressure peak increase, and the peak position moves upstream. Nevertheless, with the increase of free-stream Mach number, the upstream separation length is shortened, the pressure peak raises and the peak position shifts to the downstream. Besides, a combined parameter named momentum ratio is introduced to involve various factors, and the upstream separation length increases linearly with the momentum ratio when the momentum ratio is between 0.03 and 0.173. Finally, the influences of angle of attack on the disturbing force and moment of the upper stage are discussed. Numerical results indicate that the direction of the disturbing force is downward and the magnitude first rises and then falls with the angle of attack. The disturbing moment makes the upper stage have a tendency of subduction and increases with the angle of attack. [ABSTRACT FROM AUTHOR]

Shock wave–boundary layer interaction (SWBLI) is one of the most prevalent challenges in the field of fluid mechanics. Excessive interaction may lead to strong flow separation, which is an important factor causing the deterioration of aircraft performance. Reynolds averaged Navier–Stokes (RANS) models remain the primary methods for predicting flow separation for engineering applications. RANS simulation results typically have uncertainty owing to the drawbacks of the model structure and the variability of the parameters. Therefore, evaluation of a turbulence model based on nominal predictions may be insufficient for reflecting its actual simulation capabilities. In this study, the Bayesian method is employed to re-evaluate the simulation capabilities of three k–ω models for hypersonic SWBLI flows: the original k–ω model and two k–ω models with compressibility corrections (Wilcox's and Zeman's corrections). This is achieved by perturbing the model parameters and combining with experimental data. Bayesian model evaluation based on different datasets is performed using the embedded model error method, and the three turbulence models are compared from three aspects: posterior mean of the quantities of interests, model error, and Bayesian evidence. The results show that the Bayesian calibration can effectively improve the predictive abilities of the three models for wall pressure and wall heat flux. Different from the large differences in the nominal results, the performance of the two calibrated k–ω models with compressibility corrections are very close and superior to that of the original model. Moreover, the evaluation results of this study are also proven to be applicable to other similar flows. • The Bayesian method is employed to re-evaluate three k–ω models. • The embedded model error method is used to estimate the model error. • Turbulence models are compared from three aspects: posterior mean, model error, and Bayesian evidence. • The applicability of evaluation results is investigated. [ABSTRACT FROM AUTHOR]