Aerodynamic Throat Research Articles

Investigation of Unsteady Combustion Regimes in a Kerosene-Fueled Scramjet with Air Throttling

This paper describes an experimental study investigating unsteady combustion regimes in a kerosene-fueled scramjet. The results are obtained under inflow conditions of a 2.9 MPa stagnation pressure, 1900 K stagnation temperature, and a Mach number of 3.0. The air throttling position is 240 mm downstream of the combustor entrance, with an air throttling flow rate (ratio of air throttling mass flux to inflow mass flux) of 38% and a fuel equivalence ratio of 0.37. Combustion is relatively stable when air throttling is applied and is dominated by auto-ignition. When air throttling is turned off, the combustion becomes more unsteady and is dominated by flame propagation. At the same time, the combustion mode changes, and the frequency of the combustion mode transition is 286 Hz. Schlieren images and one-dimension analysis show that the effect of air throttling is the coupling of cold throat (aerodynamic throat) and hot throat (thermal throat). The proper orthogonal decomposition and dynamic mode decomposition analysis present that when air throttling is applied or removed, the frequencies of injector–flame feedback are almost the same, while the frequencies of shock–flame feedback exhibit considerable variation, which is caused by the location of the precombustion shock affected by air throttling.

Transient Characteristics of Fluidic Pintle Nozzle in a Solid Rocket Motor

The fluidic pintle nozzle, a new method to control the thrust of a solid rocket motor, has been proposed in recent years by combining the pintle with the aerodynamic throat (fluidic throat). The study of static characteristics has proved that it has a remarkable effect on thrust control. To study the transient characteristics of the fluidic pintle nozzle, 2D transient simulations of a fluidic pintle nozzle propulsion system were conducted, employing dynamic meshing techniques. The Reynolds-averaged Navier–Stokes equations were meticulously solved, implementing a k–ω SST turbulence model. The thrust control principle of the fluid pintle nozzle was studied, and the wave structure was summarized. The transient characteristics of the secondary flow opening, secondary flow closing, pintle moving forward (pressure rise), and pintle moving backward (pressure drop) were obtained, and the effects of the injection angle and injection port position were studied. The response process after injection can be roughly divided into three stages: pressure propagation, pressure oscillation, and equilibrium stability, with time distributions of 0.4%, 5.39%, and 94.21%, respectively. In the process of the pintle moving forward, the rate of combustion chamber pressure increases and thrust decreases gradually because of the arc wall of the nozzle throat upstream, and the process of throats moving backward is just the opposite. Compared with the condition with a maximum throat opening and no secondary flow, the thrust of the condition with a minimum throat opening and a 0.3-flow-ratio secondary flow is increased by 80.95%. Under conditions of constrained flow ratio, the injection angle of the secondary flow ostensibly exerts negligible influence on the dynamic modulation of thrust. Nevertheless, it remains evident that a reduction in throat opening accentuates the impact of reverse injection. Furthermore, the proximity of the injection port to the head of the pintle is directly proportional to the efficacy of thrust control.

Oscillations in Rectangular Supersonic Inlets with Large Internal Contraction Ratio

Increasing the internal contraction ratio (defined as the area ratio of the duct entrance and throat, denoted as ICR) can improve the operating performance of a variable geometry supersonic inlet, however, excessive ICR typically will bring the inlet into an unstarted state with oscillation. To uncover the underlying mechanism of a large ICR-induced oscillatory-unstarted phenomenon, we employed a high-speed schlieren and dynamic pressure acquisition system, and experimentally examined the flowfields of a rectangular supersonic inlet with different ICRs of 1.5, 2.0, 2.5, and 3.0 in the wind tunnel at a freestream Mach number of 2.9. The inlet stayed unstarted while the ICR ranged from 2.0 to 3.0, and the flowfield had three features. First, the inlet flowfields showed high- and broadband-frequency oscillations, whose dominant frequency and oscillatory amplitude increased with ICR. Second, the inlet flowfields revealed similar oscillatory flow structures, which contained the periodic variation of the ramp-side boundary layer separation, cowl-side flow spillage, and shock train in the internal contraction part. Third, the unstarted flow structure and oscillatory phenomenon mainly occurred upstream of the inlet geometric throat, downstream of which the main flowfield was supersonic. To characterize these three features, we proposed a self-excited oscillatory model consisting of a disturbing source, a disturbing feedback terminal, and a signal feedback loop. Practically, the disturbing source is a scale variation of the entrance separation, the disturbing feedback terminal is an aerodynamic throat near the geometric throat, and the signal feedback loop is composed of communication pathways of shock train motion, acoustic waves, and convection waves. With this model, we explained the underlying mechanism of the high- and broadband-frequency characteristics of the oscillation caused by large ICR, and the ICR’s influence on the flow oscillation.

Experimental and numerical study on reinforcement mechanism of embedded rocket on back pressure resistance of RBCC inlet at starting stage

Favorable back pressure resistance of an RBCC inlet is vital for its smooth start and stable operation in ejector-to-ramjet mode transition. As a core component of the RBCC engine, the embedded rocket can be fully used to enhance the back pressure resistance of the inlet at the starting stage. The reinforcement mechanism of an embedded rocket on the back pressure resistance of the RBCC inlet is studied through experimental and numerical methods. An RBCC inlet model fully concerning the combustion pressure in the RBCC combustor, and the expansion and aerodynamic compression of the embedded rocket jet is established, whilst a facility nozzle is deployed to replace the compression section of the inlet in the ground direct-connect tests. The results show that three typical rocket operation stages may exist impacted by both the rocket jet and the back pressure, in terms of ‘Subcritical Rocket Operation’, ‘Critical Rocket Operation’, and ‘Supercritical Rocket Operation’. Meanwhile, two different patterns of flow separation caused by the synergy of the rocket jet and back pressure are observed. In the ‘Subcritical Rocket Operation’ stage, the two separation regions are strongly influenced by the back pressure. Nevertheless, the locations of the separations and the related shock train move downstream synchronously as the rocket pressure increases. In the ‘Supercritical Rocket Operation’ stage, the increasing rocket pressure drives the cowl wall separation downstream gradually, and however, it pushes the shock train moving upstream in the isolator resulting from the aerodynamic compression of the rocket jet in the mainstream. And in the ‘Critical Rocket Operation’ stage, an optimum status is found to exist such that the shock train stands most downstream from the physical throat, at which the RBCC inlet performs at the best back pressure resistance. Therefore, an appropriate embedded rocket operation can effectively reinforce the back pressure resistance of the RBCC inlet. Since the embedded rocket is operated at relatively high pressure, the under-expanded rocket jet will compress the mainstream and form a new aerodynamic throat in the flow passage, which can: 1) increase the real internal contraction ratio of the RBCC inlet to improve the compression efficiency of the incoming flow, and 2) enlarge the aerodynamic expansion ratio of the downstream flow passage. Further, the appropriate embedded rocket operation can also effectively assist the RBCC inlet to start at a certain Mach number, but with higher combustion pressure in the ejector mode.

Aerodynamic Optimization Design of a Supersonic Compressor Rotor with High Pressure Ratio

Supersonic compressors have a high wheel speed and operational capability, which facilitate a high stage pressure ratio. However, the strong shock waves in the passage of a supersonic rotor and the interference between shock waves and boundary layers can lead to large flow loss and low efficiency. Moreover, the existing design of a high-load supersonic compressor has the problem of small stall margin. In this study, an automatic optimization method including 2D profile optimization and 3D blade optimization is proposed to achieve a high efficiency at the design point of a supersonic compressor rotor under the premise of reaching the desired mass flow rate and total pressure ratio. According to the analysis of flow near the stall point of the supersonic compressor rotor, the mechanism responsible for rotor tip stall is established, that is, the aerodynamic throat appeared inside the flow passage, reducing the ability of the blade tip to withstand back pressure, and the low-speed areas caused by the tip-leakage-vortex breakage and boundary layer separation reduced the flow capacity of the blade tip. Based on the reasons for rotor stall, three methods are proposed to improve the stall margin, which include increasing the exit radius of the upper meridian, forward sweep of the blade tip, and increasing the chord length of the blade tip. The above method is used to design a supersonic rotor with a total pressure ratio of 2.8, which exhibits an efficiency of 0.902 at the design point and a stall margin of 18.11%.

The effect of sweep on flowfields of a highly loaded transonic rotor

A new view of internal blade loading match is employed in this paper to reveal the essential relationship between sweep and complicated flow phenomenon in compressors. A forward swept rotor named as SF and an aft swept rotor named as SA, are designed originating from NASA Rotor 37. The numerical results indicate the isentropic efficiency and total pressure ratio of the new rotor with forward sweep are a little larger than rotor Baseline at peak efficiency point, and its stall margin has an over 10% increase as well. By contrast, the opposite change appears in the aft swept rotor SA. In addition, the choking massflow of swept rotors changes clearly as a result of the spatial change of aerodynamic throat and the forward sweep produces a larger throat area and choking massflow. A noted phenomenon is shown that both forward and aft sweep benefit a certain range of blade span for flow characteristic improvement, and the sweep also affects the inflow condition, 3-D shock structure and internal loading distributions of swept rotors. Both the new radial equilibrium in flowfield and transportation of low-energy fluids in boundary layer induced by sweep appear conductive to reduce shock wave and flow blockage in tip region of forward swept rotor SF. Meanwhile, the strength of leakage flow is also suppressed in the forward swept rotor SF because of the decreased blade loading near leading edge in tip region. In fact, all the change in swept rotors tends to be a new characteristic match of blade airfoil at different span, and the re-matching of flow characteristic for blade airfoil determines the stable operation range and the overall performance of swept rotors at design point directly.

Study on Flame Stabilization in a Dual-Mode Combustor Using Optical Measurements

Flame stabilization in a dual-mode scramjet combustor was studied using simultaneous detection of flowfield and reaction zone. The instantaneous reactive flowfield was clearly visualized using an improved pulsed schlieren system, whereas the reaction zone was marked by chemiluminescence. Experiments were performed in a Mach 2.5 model combustor with the total temperature of and the total pressure of 1.0 MPa. Ethylene was used as fuel with equivalence ratios varied from 0.258 to 0.411. Four typical flame stabilized locations and corresponding flowfields are presented. When no choke occurs, the flame is stabilized in the cavity or the shear layer. Schlieren images show that the flow in the reaction zone of the latter case is supersonic. When the combustor is choked at the injection location, the flame oscillates between the shear layer and the jet wake, and the flow in the reaction zone becomes subsonic. When the flame is stabilized in the jet wake, autoignition can occur due to the compression of precombustion shock train. The short-lived aerodynamic throat formed near the injection location may be the possible cause for triggering the unsteady flame oscillation. The effects of heat release on flame stabilization modes are discussed.

Convergent/divergent segmented exhaust nozzle

A segmented exhaust nozzle for attenuating noise from a turbofan jet engine without adversely impacting the operability or operability limit related performance of the engine. The exhaust nozzle includes spaced apart fan nozzle inner and outer walls which form an annular exhaust gas flow path therebetween. The fan nozzle outer wall is segmented at the downstream end. The outer wall curves inwardly towards the inner wall and then turns back away from the inner wall to form an arcuate protrusion that extends into the exhaust gas flow path forming an aerodynamic throat. Through the segmented portion of the nozzle, the outer wall then continues to curve away from the inner wall before again curving back towards the inner wall at a nozzle exit station. The nozzle exit effective area is approximately equal in cross sectional area to a conventional exhaust nozzle exit area. The inwardly curving and then segmented outwardly curving portion of the exhaust nozzle forms a geometric influction that serves to reduce noise without negatively affecting engine operability or operability limit related performance.

Steady and Unsteady Flow Phenomena in a Channel Diffuser of a Centrifugal Compressor

The aim of this paper is to understand the time averaged pressure distributions and unsteady pressure patterns in a channel diffuser of a centrifugal compressor. Pressure distributions from the impeller exit to the channel diffuser exit are measured and discussed for various flow conditions. And unsteady pressure signals from six fast-response sensors in the channel diffuser are analyzed by decomposition method. The strong non-uniformity in the pressure distribution is obtained over the diffuser shroud wall caused by the impeller-diffuser interaction. As the flow rate increases, flow separation near the throat, due to large incidence angle, increases aerodynamic blockage and reduces the aerodynamic flow area downstream. Thus the minimum pressure location occurs downstream of the geometric throat, and it is named as the aerodynamic throat. And at choke condition, normal shock occurs downstream of this aerodynamic throat. The variation in the location of the aerodynamic throat is discussed

A Study on Pressure Distributions in a Centrifugal Compressor Channel Diffuser

Time averaged pressure distributions in a high-speed centrifugal compressor channel diffuser at design and off-design flow rates are investigated. Pressure distributions from the impeller exit to the channel diffuser exit are measured for various flow rates from choke to near surge condition, and the effects of operating condition are discussed. The strong non-uniformity in the pressure distribution is obtained over the vaneless space and semi-vaneless space caused by the impeller-diffuser interaction. As the flow rate increases, flow separation near the throat, due to large incidence angle at the vane leading edge, increases aerodynamic blockage and reduces the aerodynamic flow area downstream. Thus the minimum pressure location occurs downstream of the geometric throat, and it is named as the aerodynamic throat. And at choke condition, normal shock occurs downstream of this aerodynamic throat. The variation in the location of the aerodynamic throat is discussed.

Flowfield measurements inside a boundary-layer bleed slot

An experimental investigation was conducted to determine the flowfield inside a bleed slot used to control an oblique shock-wave and turbulent boundary-layer interaction. The slot was oriented normal to the primary flow direction and had a width of 1.0 cm (primary flow direction) and a length of 2.54 cm and spanned 16.5 cm. The approach boundary layer upstream of the interaction was nominally 3.0 cm thick. Two operating conditions were studied: M = 1.98 with a shock generator deflection angle of 6 deg and M = 2.46 with a shock generator deflection angle of 8 deg. Measurements include surface and flowfield static pressure, pitot pressure, and total mass flow through the slot. The results show that despite an initially two-dimensional interaction for the zero-bleed-flow case, the slot does not remove mass uniformly in the spanwise direction. Inside the slot, the flow is characterized bv two separation regions, which significantly reduce the effective flow area. The upper separation region acts as an aerodynamic throat, resulting in supersonic flow through much of the slot.

Multiple ducted streams with a periodic or a steady supersonic driver flow

AN experimental investigation of the operating characteristics of a two-dimensional planar, constant area, supersonic-subsonic ejector with a steady or a periodically pulsed continuous driver flow is presented. Improvements in ejector pressure recovery and entrainment capabilities resulted from the periodic driver for a limited range of operating conditions. Contents The purpose of this study was to determine if periodic pulsation of the driver flow in an ejector would lead to improvements in ejector performance as a result of enhancing or exciting modes of flow interaction beyond those normally occurring in the ejector mixing duct. In particular, this investigation was concerned with pressure recovery applications in which the primary flow must pump the low pressure secondary to an appreciably higher pressure level. The approach taken to induce flow interaction was to periodically pulse the ejector driver flow with an efficient large scale fluidic oscillator developed for this purpose. This is similar to wave energy exchanger type ejectors13 in which slugs of primary flow are discharged periodically into the mixing duct. In the current case, however, the driver flow is continuous. A simple, one-dimensional inviscid control volume analysis, following that of Fabri and Siestrunk,4'5 was used to predict ejector operating characteristics; the descriptive features of the analysis are useful for the present discussion. Three regimes of ejector operation are considered. The supersonic regime, SR, occurs when the primary flow expands against the secondary after entering the ejector mixing duct and the secondary flow chokes at an aerodynamic throat in the mixing duct; this is the situation depicted in Fig. 1. The saturated supersonic regime, SSR, occurs when the secondary static pressure equals or exceeds that of the primary flow at the ejector inlet and the secondary flow chokes at the inlet. The mixed regime, MR, occurs when the back pressure at the ejector exit plane is too high to allow choking of the secondary flow either in the mixing duct or at the inlet.

Heavy Fuels Ash Deposit Formation and Removal in Water-Cooled High-Temperature Gas Turbines

Experimental data obtained in heavy fuels operation of a gas turbine simulator with a water-cooled, transonic turbine nozzle cascade are presented. The ash fouling is characterized by the rate of decrease of the aerodynamic throat area. Particular attention is given to the cleanability of the ash deposits. A simple heat transfer analysis was performed to assist in evaluating the data. The rate of ash fouling in the water-cooled nozzle was found to be of the same order of magntude as for conventional air-cooled designs. Cleanability, both on and off-line, was found to be significantly enhanced, thus making the water-cooled gas turbine an attractive alternative for heavy fuels applications.

Experimental investigation of a novel vtol thrust vectoring nozzle

A two-dimensional corner-expansion thrust vectoring nozzle, in which thrust vectoring is achieved by rotation of a single vane, is proposed for high supersonic V/STOL aircraft. A -fa scale model of the configuration has been statically tested with most favorable results. A thrust coefficient in excess of 0.97 was demonstrated in the VTOL mode and throughout transi- tion at nozzle pressure ratios typical of turbojet engines considered for high supersonic V/ STOL applications. In addition, a thrust coefficient in excess of 0.96 was attained in the horizontal cruise mode down to a nozzle pressure ratio of 3.5. This is particularly significant when considering that the design pressure ratio of the nozzle was 21. Effective thrust vectoring was also demonstrated, with a 1:1 correspondence between vane mechanical deflection and thrust vector direction. A jet pumping effect was found to exist at very low-pressure ratios at a slightly deflected position of the thrust vectoring vane, and an alternating normal compon- ent of the total thrust vector was found to exist at low-pressure ratios in the horizontal cruise mode. studies have generally shown that thrust vectoring is a favorable scheme for high supersonic V/STOL interceptor missions, thus establishing the need for developing a simple, highly efficient, and reliable thrust vectoring nozzle; in addition to providing the basic V/STOL thrust vectoring and low-pressure ratio cruise performance for subsonic and low supersonic missions, an efficient nozzle is required for the high supersonic cruise condition in which the nozzle pressure ratio may be as high as 30. A simple and efficient thrust vectoring nozzle, fulfilling all the requirements of a high supersonic V/STOL mission, has been conceived and tested at Nor air with favorable results. The detailed results of this test are documented in Ref. 2. The basic features of the configuration, which conveniently lends itself to underbody installations, are shown in Figs. 1 and 2. In the horizontal flight position, the nozzle geometry is that of an isentropic wedge in which a two-dimensiona l free- jet corner-expansion to the local ambient pressure occurs at the nozzle lip. The upper boundary of the nozzle is denned analytically by a Prandtl-Meyer stream line at the nozzle design pressure ratio. Ideally, an isentropic expansion occurs at the design pressure ratio giving a horizontal free- jet boundary downstream of the nozzle lip, with uniform, parallel, fully expanded flow at the axial location defined by the intercept of the nozzle upper boundary with the design Mach number line emanating from the nozzle lip. At off- design pressure ratios, nozzle losses due to overexpansion and underexpansion of the exhaust gases are minimized by the ability of the lower jet boundary to adjust to the external local pressure as for a conventional plug nozzle. Side plates are provided to contain the jet in the transverse direction in the region of the flow field where jet pressures are greater than ambient. The side plates eliminate end-effect losses by maintaining a constant and maximum pressure across the width of the upper boundary, thus insuring the maximum value of axial thrust. Variable nozzle throat area in the aft position, if desired, can be conveniently provided by rotation of the nozzle lip. In the VTOL and transition modes of operation, an efficient free-jet aerodynamic-throat nozzle is formed by rotation of the thrust vectoring vane and associated extension. In these modes of operation, the flow expands around the sharp corner at the terminal point of the nozzle lip with a free-jet boundary and aerodynamic throat prevailing downstream of the nozzle lip. Ideally, the flow is expanded supersonically and uni- formly at the vane trailing edge. In principle, the internal aerodynamic operation of the nozzle in the VTOL and transi- tion modes is much like that of the horizontal cruise configura- tion, with the essential difference being that the locus of the sonic line in the former modes occurs downstream of the minimum physical internal cross section of the nozzle and, of course, the design pressure ratio is an order of magnitude lower.

Automatic setting of the flexible walls of a large wind tunnel

To develop supersonic air flow in a wind tunnel it is usual to provide a nicely contoured constriction in the air passage ahead of the model test region, and it is necessary to alter the cross-section of this ‘throat’ to cater for different air speeds. In the tunnel under consideration, the constriction is formed by two flexible steel plates, sliding between side walls and constrained to the required adjustable profile by a series of hydraulically-actuated screw jacks.The paper deals with the electrical control system, for storing the necessary jack-setting data on punched tape, and releasing it to govern, simultaneously, the motion of every screw jack with the required degree of precision, so that the shape of the aerodynamic throat can be adjusted progressively throughout the air-speed range of the tunnel.In developing the system, a number of pieces of electro-mechanical equipment had to be devised. Some of this new apparatus is briefly described, and its operation is indicated.