Volume 38, Issue 3

The interaction of Mach 0.5, 0.8, and 1.0, parallel, twin circular jets issuing from orifices with center-to-center spacing S / D , where S is the center-to-center distance and D is orifice diameter, 2, 4 and 6 has been investigated experimentally. The characteristics of twinjets are analyzed based on the centerline Mach number decay, exit Mach number and ratio of orifice spacing. As the spacing between the orifice increases, the maximum Mach number point of the combined jet moves downstream. For the Mach numbers studied it is found that as the S / D increases the effect of the counter-rotating vortices on jet mixing decreases. The rate of the twinjet interaction also decreases with S / D increase. As the jets propagate downstream their center-to-center distance decreases continuously and the jets merge to become single jet, for all S / D studied.

Centrifugal compressors have been widely used in light-duty turboshaft engines and automotive turbochargers. One of the essential requirements for centrifugal compressors in these applications is to have adequate stable flow range. As the pressure ratio increases, the stable flow range drops dramatically, and this highlights the importance of stability improvement for centrifugal compressors. This paper first gives a brief review on the measures available for extending the stable flow range, and then focuses on the realization of self-recirculating casing treatment for a centrifugal compressor. The design principles of the casing treatment were introduced. Then the simulations were conducted to design and optimize the geometry of the casing treatment, and finally an optimum case was selected for experimental validation. The experiment results indicate that the designed casing treatment improve the stable flow range by about 5 percent at design speed, while keeping the efficiency almost identical.

A streamline curvature method based integrated throughflow analysis approach is newly developed to deal with component matching problems of variable cycle compression systems. The construction of variable cycle compression system is modularly modelled in the procedure. Splitting and confluent flow are elaborately disposed. A numerical method based on the “streamline floating” character of streamline curvature method is developed to model the function of forward variable area bypass injector. Moreover, extensive models used in the throughflow calculations, including minimum loss incidence, deviation and loss models were assessed, selected and modified. Finally, code validations were conducted on three representative traditional compressors, i. e. NASA rotor 67, NASA stage 37 and a custom-designed low-speed repeating four-stage compressor. Both the predicted overall characteristics and spanwise profiles agree reasonably well with the experimental data. The validated procedure was finally used to sketch the performance maps of a double bypass compression system under two different control rules, i. e. the first bypass throttling and the second bypass throttling. The results show some aspects of the difficulties and complications in operating a variable cycle compression system, and meanwhile, demonstrate the superiority of the newly developed integrated throughflow method.

Computational investigations on the effect of Gurney flaps in a NACA 0010 cascade with a 90 o blade angle were carried out using STARCCM+ code. The flaps were positioned at the trailing edge of the blades and were aligned at 45 o and 90 o to the blade chord. The variations in the lift, drag, aerodynamic efficiency and static pressure distributions were recorded on the blades and were compared with the baseline case. Significant increase in the lift coefficient was observed due to the Gurney flaps accompanied by a small drag penalty. This is achieved by increasing the effective camber of the blade, thereby increasing the pressure difference between the pressure and suction surfaces. The Gurney flaps increased the aerodynamic efficiency of the blade at lower angles of attack, whereas at higher angles a small drop in the value was observed.

In order to describe how the variant BPR affects the flow performance, a qusi-2D fan model was proposed. In the model, the stage cumulative method was applied to compute axial flow and the simplified radial equilibrium equation was applied to compute radial flow, then a dual-level iteration program was proposed for the nominal performance calculation. To compute the non-nominal flow characteristics on non-nominal BPR, a triple-level iteration program with the compensation factor was proposed to obtain the radial profiles. On that basis, Blade angles were designed for this model on the design point. With the inflow deflection factor designed as 0.8 and the loss factor along blade as 0.95, the pressure ratio error is only 0.21 % and the isentropic efficiency error is 0.06 %. Simulation showed that the variant-BPR profiles and the overall characteristics of flow at fan exit or for bypass & core ducts could be well computed with much less time-consuming.

The control of co-flowing jets by varying lip thickness has been studied experimentally. Lip thickness is defined as the thickness of primary nozzle wall separating primary jet and secondary jet at the co-flowing nozzle exit. Co-flowing jets from a primary nozzle of diameter 10 mm (1.0 D p ) and a secondary duct with lip thickness (LT) 0.2 D p , 1.0 D p and 1.5 D p at Mach numbers 0.6, 0.8 and 1.0 have been studied. Jet centreline total pressure decay, static pressure variation and jet mixing behaviour were analysed. The results show that the mixing of the co-flowing jet with substantial values of lip thickness is superior to the co-flowing jets with comparatively lower values of lip thickness. Co-flowing jets with lip thickness 1.0 D p and 1.5 D p experience a significantly higher mixing than the lip thickness 0.2 D p jet, for all Mach numbers analyzed in the present study. Moreover, in the case of correctly expanded jets, the local static pressure is assumed to be equal to atmospheric pressure. This assumption becomes invalid for co-flowing jets with substantial lip thickness. The centerline static pressure varies sinusoidally above and below atmospheric pressure by a maximum of 11 %, which is due to wake dominance.

Trailing edge slot film cooling is a widely used method for protecting the trailing edge of turbine blades from hot gas impingement. The structures that separate the slots, known as “lands,” come in a variety of configurations. This paper presents the effects of the trailing edge with different lands on the film cooling performance. Experimental studies are conducted on the film cooling effectiveness and Nusselt number with different lands. Four trailing edge configurations, including the straight lands, the beveling lands, the fillet lands and the tapered lands are considered under four blowing ratios (0.5, 0.7, 1.0 and 1.5). The Reynolds numbers of mainstream is fixed as 375,000. Film cooling effectiveness and Nusselt number performances are measured by transient liquid crystal measurement technique. Reynolds-averaged Navier-Stokes (RANS) simulation with realizable k-ε turbulence model and enhanced wall functions are performed using a commercial code Fluent. In each case, the slot height is kept constant. It is shown that the beveling lands, the fillet lands and the tapered lands have higher cooling effectiveness and lower Nusselt number compared with the straight lands. Under higher blowing ratios, the trailing edges of all four lands have higher cooling effectiveness and higher Nusselt number.

As per few investigation in installed performance for variable cycle engines, an analysis scheme is proposed on the basis of integrating variable cycle engine and supersonic inlet system model. An integrated mathematical model, containing the inlet and the variable cycle engine is built, realizing the simulation of influences on the installation performance by varying geometry components. The impact on engine performance of variable geometric regulation was analyzed and concluded respectively. The experimental results show that the overflow resistance of the variable cycle engine with variable geometry optimization is reduced at subsonic cruise stage, and the installed fuel consumption is reduced, which significantly improves the installation performance.

Control of Mach 1.8 circular jet with slanted perforated tabs is studied experimentally. Two sets of perforated tabs were used for this study. The perforation angles were 0° and 30° with respect to axis of the nozzle. The blockage areas of the tabs were 5 %. The mixing enhancements caused by these tabs were studied in the presence of adverse and favorable pressure gradients, corresponding to nozzle pressure ratio (NPR) of 4, 5.74 and 8. For Mach number 1.8, jet NPR 4 corresponds to 30 % adverse pressure gradients and NPR 8 corresponds to 39.37 % favorable pressure gradients. The pressure decay characteristics and shadowgraph images of perforated tabs at different NPR were compared. There is 45 % and 65 % reduction in jet core length were observed for the 0° and 30° perforated tabs respectively in both pitot and shadowgraph experiments in comparison to uncontrolled jet.

Characteristics of full coverage film cooling of an adiabatic flat plate are studied for opposite injection of coolant at different angles. Two in-line adjacent rows of cooling holes injecting in opposite directions are considered in this study. The cooling performance is compared with the configurations having forward and reverse injecting holes at similar injection angles. The holes are arranged in an array of 20 rows with equal spacing both span-wise and stream-wise. Computational analyses are carried out over a wide range of velocity ratios ( VR ) of practical importance ranging from 0.5 to 2.0 at density ratio of about 1.0. Injection angle and velocity ratio are found to have strong influence on film cooling effectiveness of opposite injection. At low velocity ratio of VR =0.5, film cooling performance of opposite injection at 45° is found better than at other angles, i. e. 30° and 60°. At higher velocity ratios, injection at 30° is found superior. Film cooling effectiveness becomes insensitive to velocity ratios at higher range for 45° and 60° injections. Evolution of effusion film layer and interaction between coolant and primary flow is also studied in this paper.