Geometric Throat Research Articles

An extension of the compound flow theory with friction between the streams and at the wall

Compound flows consist of two or more parallel compressible streams in a duct and their theoretical treatment has gained attention for the analysis and modelling of ejectors. Recent works have shown that these flows can experience choking upstream of the geometric throat. While it is well known that friction can push the sonic section downstream of the throat, no mechanism has been identified yet to explain its displacement in the opposite direction. This study extends the existing compound flow theory and proposes a one-dimensional (1-D) model, including friction between the streams and the duct walls. The model captures the upstream and downstream displacements of the sonic section. Through an analytical investigation of the singularity at the sonic section, it is demonstrated that friction between the streams is the primary driver of upstream displacement. The 1-D formulation is validated against axisymmetric Reynolds averaged Navier–Stokes simulations of a compound nozzle for various inlet pressures and geometries. The effect of friction is investigated using an inviscid simulation for the isentropic case and viscous simulations with both slip and no-slip conditions at the wall. The proposed extension accurately captures the displacement of the sonic section, offering a new tool for in-depth analysis and modelling of internal compound flows.

Combustion organization strategies for a variable geometry RBCC combustor under low total temperature conditions

A rocket-based combined cycle (RBCC) engine experiences a low Mach number phase during flight operations. Through combustor geometry adjustment technology, the engine can combust more efficiently under low-temperature inflow conditions during this phase, thereby improving the engine efficiency. In this paper, we investigate the fuel blending and combustion processes in rocket ramjet and pure ramjet modes under Ma 2.5 low-temperature inflow conditions, analyze the fuel combustion and flow and engine performance changes under different operating modes, and propose a new method for centralized heat release combustion organization under low-temperature inflow conditions. The method is based on the RBCC engine combustor with a geometry throat and relies on both direct-connect experiments and numerical calculations. The results show that (1) when the primary rocket operates, its jet reacts with the inflow in the shear layer and ignites the kerosene fuel injected by the pylons, forming a high-temperature zone downstream of the fuel pylons. When the primary rocket stops working, this flame is stabilized in the low-speed reflux region at the exit of the rocket by injecting a small equivalent ratio of fuel in the isolator to mix with the air. This approach can be a good substitute for the rocket jet to stabilize the flame. (2) When the primary rocket operates, the secondary fuel is mainly concentrated in the area from the fuel pylons to the geometric throat. After the rocket is turned off, the interval of heat release from combustion is also concentrated in the combustor from the fuel pylons to the geometric throat. (3) After the primary rocket stops working, the performance of the combustor improves significantly.

Investigation of the influence mechanism of geometric throats in variable-geometry rocket-based combined-cycle combustors

To explore the reaction mechanism a geometric throat and its influence on the performance of the combustor, this paper studies the working process of a rocket-based combined-cycle (RBCC) variable geometry combustor with geometric throats under Ma4 inflow conditions by both numerical simulation and ground direct-connect testing. The main research contents and conclusions are as follows: (1) The geometric throat can accurately control the choking position. The fuel in the variable geometry combustor can complete the concentrated heat release and bring higher combustor pressure. The thrust of the geometric throat combustor is 8.7% higher than that of the thermal throat combustor. Therefore, the geometric throat combustor provides a better combustor performance than the thermal throat combustor. (2) When the throat area of the Ma4 configuration is adopted, the thrust and specific impulse in the combustor are at their the highest, reaching 6.0% and 0.4% higher than those of the Ma5 and Ma3 throat configuration combustors, respectively. This shows that a certain amount of combustor heat release needs to match the corresponding geometric throat area to ensure the optimal performance of the combustor. (3) In the Ma4 combustor configuration, after closing the primary rocket, the thrust in the combustor decreases by approximately 10.8%, but its specific impulse performance greatly improved by 39.3%. Therefore, the geometric throat can be used in the RBCC mode transition process in combination with the secondary fuel injection strategy to avoid the thrust trap in the mode transition process.

Investigation on geometric throat matching characteristics of variable geometry rocket-based combined cycle combustor

One of the most effective ways to realize multi-mode and high-performance matching operation of rocket-based combined cycle (RBCC) engine is to adopt a variable geometry combustor. In this study, direct-connect experiments and three-dimensional numerical simulations were used to investigate the matching characteristics between the isolator and the geometric throat of the variable geometry rocket-based combined cycle combustor under Ma 6 condition. The influences of equivalence ratio and geometric throat height variations on the combustor matching characteristics were obtained. The study results show that: (1) The pressure in the combustor and the outlet pressure of the isolator increase gradually with the increase of equivalence ratio, which causes the pre-combustion shock train position in the isolator to move forward gradually. Meanwhile, the combined injection of secondary fuel with the isolator and the pylon is beneficial to the improvement of combustor pressure, thus further improving the engine performance. (2) Different geometric throat heights have a substantial impact on the isolator and combustor matching when the equivalent ratio of the combustor is fixed. Under the conditions of Ma 6 operating at an altitude of 24 km inflow, the starting position of the pre-combustion shock train in the isolator is precisely at the isolator entrance, and the corresponding geometric throat height is 2.10H. The isolator and the combustor can complete matching operation at this point which the geometric throat height is the minimal throat height for Equivalent ratio = 1.0. (3) The reduction in geometric throat height can result in an increase in combustor pressure but a loss in combustor heat release and combustion efficiency of fuel combustion, which are reduced by 21.1% and 7.2% respectively. (4) The heat release and combustion efficiency in the combustor will decrease with a decrease in equivalent ratio at the minimal throat height condition where the isolator and the combustor can complete matching operation. This happens when the equivalent ratio and the geometric throat height change synergistically.

Nozzle Design for Rotating Detonation Engine

This work proposes a nozzle design procedure that uses an area-variational source term in the two-dimensional numerical simulations for rotating detonation engines. The derivation of this source term is carried out in detail, and the influences of area ratios, total lengths, and curve transitions on the aerodynamic performance are discussed. For a nozzle with no curve transition at critical cross-sections, the combination of contraction ratio and expansion ratio corresponds to the maximum axial thrust and maximum specific impulse at and , which exceed those of the diverging and converging nozzles by 11.3 and 27.6%, respectively. The extension of the total length is not recommended to attenuate fluctuations due to the thrust performance penalties. Concerning the curve transitions of the diverging part, the parabolic profile subp2 shows a 2.22% increase in thrust and specific impulse compared with four other types. The positive influence of a geometrical throat on the thermodynamic efficiency is demonstrated by the fluid particle tracing technique and the pressure gain analysis. Moreover, the feasibility of adding the area-variational source term is validated by three-dimensional simulations.

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.

A Parametric Study on the Fluid Dynamics and Performance Characteristic of Micronozzle Flows

Abstract This study investigates the fluid dynamics and performance characteristics in micronozzle flows with changes in various geometric parameters using Navier–Stokes simulation based on slip wall boundary conditions. The various geometric parameters considered for the study are (1) area ratio with fixed throat dimension and (2) the semidivergence angle variation with no change in area ratio. The simulation results show that the flow choking for micronozzle happens not at the geometric throat; rather pushed downstream to the divergent channel of the nozzle. This is due to the thick boundary layer growth, which reduces the effective flow area and shifts the minimum allowable flow area downstream to the throat. The distance to which the choking point shifts downstream to the throat reduces with Maxwell's slip wall conditions compared to the conventional no-slip wall condition. The downstream movement of the choking point from the throat reduces with an increase in area ratio and with increase in divergence angle with fixed area ratio. This is due to the fact that the increase in area ratio and divergence angle increases the nozzle height at any particular section in the divergent portion of the nozzle. As a result of this, the boundary layer profile also moves upward and the restriction of potential core by the thick boundary layer reduces, which in turn leads to an increase in the effective minimum flow area downstream to the throat.

Development of Choked Flow in Variable Nozzle Radial Turbines

In commonly applied one-dimensional choking models for radial turbines, choked flow is assumed to appear in the geometrical throat of each stator and rotor. Coupled and complex three-dimensional effects are not considered. In order to analyze the internal aerodynamic in a radial turbine at off design conditions and before carrying out experimental tests, which in the case of automotive turbocharger are limited by their compact size, computational fluid dynamics (CFD) simulations stand out as a useful tool. This paper presents the study of a variable geometry turbine (VGT) of a commercial turbocharger at off design conditions reaching choked flow, analyzing the presence of this limiting conditions in the stator and rotor under different operation points and VGT positions. Reynolds-averaged Navier-Stokes (RANS) and unsteady RANS simulation have been performed to obtain the flow structures in stator and rotor. The results reveal that the choked effective area mostly depends on the stator vane position and pressure ratio. For the closed VGT position a standing shock wave appears on the stator suction side and expands through the vaneless space. For the opened VGT position the flow is choked at the rotor outlet. However, the evolution of the choked area highly depends on the rotational speed and the secondary flow. A strong interaction with the tip leakage vortex has been identified.

Civil turbofan engine exhaust aerodynamics: Impact of fan exit flow characteristics

It is envisaged that future civil aero-engines will operate with greater bypass ratios compared to contemporary configurations to lower specific thrust and improve propulsive efficiency. This trend is likely to be accompanied with the implementation of a shorter nacelle and bypass duct for larger engines. However, a short bypass duct may result in an aerodynamic coupling between the exit flow conditions of the fan Outlet Guide Vanes (OGVs) and the exhaust system. Thus, it is imperative that the design of the exhaust is carried out in combination with the fan exit profile. A parabolic definition is used to parameterise and control the circumferentially-averaged radial profiles of stagnation pressure and temperature at the fan OGV exit. The developed formulation is coupled with a parametric exhaust design approach, an automatic computational mesh generator, and a compressible flow solution method. A global optimisation strategy is devised comprising methods for Design of Experiment (DOE), Response Surface Modelling (RSM), and genetic optimisation.A combined Design Space Exploration (DSE) comprising both geometric, as well as fan exit profile variables, is performed to optimise the exhaust geometry in conjunction with the fan exit profile. The developed approach is used to derive optimum exhaust geometries for a tip, mid, and hub-biased fan blade loading distribution. It is shown that the proposed formulation can ameliorate adverse transonic flow characteristics on the core after-body due to a non-uniform bypass inflow. The hub-loaded profile was found to be most penalising in terms of exhaust performance compared to the mid and tip-loaded variants. It is demonstrated that the combined fan exit profile and exhaust geometry optimisation offers significant performance improvement compared to the fixed inflow cases. The predicted performance benefits can reach up to 0.19% in terms of exhaust velocity coefficient, depending on fan loading characteristics. A notable improvement is also noted in terms of bypass nozzle discharge coefficient. This suggests that the combined optimisation can lead to an exhaust design that can satisfy the engine mass-flow rate demand with a reduced geometric throat area, thus potentially offering further exhaust size and weight benefits.

A Study on Matching Between Centrifugal Compressor Impeller and Low Solidity Diffuser and Its Extension to Vaneless Diffuser

A centrifugal compressor requires a wide operating range as well as a high efficiency. At high pressure ratios, the impeller discharge velocity becomes transonic and effective pressure recovery in a vaned or vaneless diffuser is necessary. At high pressure ratios, a vaned diffuser is used as it has high pressure recovery, but may have a narrow operating range. At low flow, diffuser stall may trigger surge. At high flow, choking in the throat of the vanes may limit the maximum flow rate. A low solidity diffuser allows a good pressure recovery because it has vanes to guide the flow and a wide operating range as there is no geometrical throat to limit the maximum flow. In experimental studies at a pressure ratio around 4:1, the author has replaced vaned diffusers with a range of low solidity diffusers to try to broaden the operating range. The test results showed that the low solidity diffuser also chokes. In this paper, a virtual throat is defined and its existence is confirmed by flow visualization and pressure measurements. A method to select low solidity diffusers is proposed based on test data and the fundamental nature of the flow. The extension of the proposed method to the selection of a vaneless diffuser is examined and a design approach for a vaneless diffuser system to minimize surge flow rate without limiting the attainable maximum flow rate is proposed.

Propulsive Performance and Heating Environment of Rotating Detonation Engine with Various Nozzles

Geometric throats are commonly applied to rocket combustors to increase pressure and specific impulse. This paper presents the results from thrust measurements of an ethylene/gas-oxygen rotating detonation engine with various throat geometries in a vacuum chamber to simulate varied backpressure conditions in a range of 1.1–104 kPa. For the throatless case, the detonation channel area was regarded to be equivalent the throat area, and three throat-contraction ratios were tested: 1, 2.5, and 8. Results revealed that combustor pressure was approximately proportional to equivalent throat mass flux for all test cases. Specific impulse was measured for a wide range of pressure ratios, defined as the ratio of the combustor pressure to the backpressure in the vacuum chamber. The rotating detonation engine could achieve almost the same level of optimum specific impulse for each backpressure, whether or not flow was squeezed by a geometric throat. In addition, heat-flux measurements using heat-resistant material are summarized. Temporally and spatially averaged heat flux in the engine were roughly proportional to channel mass flux. Heat-resistant material wall compatibility with two injector shapes of doublet and triplet injection is also discussed.

Significance of gas-liquid interfaces for two-phase flows in micro-channels

Investigation was designed to clarify the difference in resistance for single-phase and two-phase flows in constricted capillaries, to define the effective pore throat, and at what pore size of a micro-channel the capillary resistance to gas-liquid interfaces starts taking effect. The experimental results indicate that the pressure drop profile for two-phase flows is significantly different from that for single-phase flows in constricted capillaries. For the same capillary, the resistance to a single-phase flow keeps constant after initial increase due to the fluid acceleration, but for two-phase flows, the resistance increases sharply when the two-phase interface enters the channel with a diameter smaller than a certain size. We defined this ‘certain size’ as the ‘effective pore throat’. For a flow channel with a diameter larger than the effective pore throat, the capillary resistance to the interface is almost zero. When the flow channel has a size less than the effective pore throat, capillary force to two-phase interfaces takes effect, and the resistance to two-phase flows increases suddenly. The ‘effective pore throat’ is a critical point to determine whether the capillary resistance to the interface takes effect in a confined flow channel. It is between 150 and 650 µm depending on capillary tip size and is very different from the geometrical throat of a channel. To predict pressure drops for a two-phase flow in constricted capillaries, a new equation is derived based on Darcy-Weisbach equation to calculate the frictional pressure drop in constricted capillaries. Young-Laplace equation is used to calculate the capillary pressure drop. The results show that after the interface enters the channel with a diameter less than the ‘effective pore throat’, the combination of our newly derived equation and Young-Laplace equation can predict the pressure drop with a deviation of ±20%. However, theories in literature cannot explain the sudden increase in the pressure drop and simulations cannot predict the effective pore throat.

Numerical Investigation of the Influence of Circumferential Flow Inhomogeinity on the Power Loading of HPT Blades

One of the methods used for the reduction of dynamic stresses in turbine blades when operating conditions are close to resonance conditions can be the reduction of external exciting forces. This scientific paper gives the data of numerical investigation of the influence of circumferential inhomogeneity before the turbine wheel on the level of nonstationary aerodynamic forces applied to turbine blades. With regard to the numerical flow simulation the full temperature field inhomogeneity was specified at the turbine inlet including the film deposition of the nozzle diaphragm. Nonstationary flows in the high pressure turbine stage were calculated for the three geometrical options of nozzle diaphragm: 1 – the initial geometry has the same area of geometric throat in each blade channel; 2 – the geometry obtained by the interlacing the two types of sectors; each includes three channels that have the enlarged area а 2 or the reduced area а 1 of throat; 3 – the geometry obtained by the interlacing of channels that have the enlarged area а 2 or the reduced area а 1 of throat; Using the Fourier –analysis data for nonstationary forces the spectral analysis of the dynamic component of these forces was done. A decrease in the level of nonstationary aerodynamic forces of z∙fn frequency was obtained for the turbine blade, where fn is the rotor speed, z is the number of blades in the nozzle diaphragm.

Acoustically Induced Shock Oscillations in a Low-Boom Inlet

Experimental unsteady centerbody surface pressures measured in a low-boom inlet have been analyzed and compared with an unsteady computational flow approach. The experimental dataset was gathered at the supersonic wind tunnel at the NASA John H. Glenn Research Center in 2010. The axisymmetric external compression inlet considered herein featured a relaxed isentropic centerbody compression spike followed by a short subsonic diffuser to the aerodynamic interface plane. The axisymmetric inlet computational domain comprised a 10 deg sector, starting with the freestream inflow region, and included both the internal flowpath up to the mass flow plug and the external flow past the sharp-edged cowl for external flow. The selected inlet test conditions were based on a 1.67 freestream Mach number at a zero angle of attack with a near-design spillage rate of approximately 4%. Both experiments and simulations revealed temporal shifts between pressure peaks at different streamwise locations, indicating upstream-running compression waves that moved at acoustic speeds. These waves became amplified near the geometric throat and produced streamwise oscillations of the external normal shock. The simulations also revealed a second smaller shock on the centerbody at the geometric throat, for which the complex dynamics suggested an opportunity for further study.

Slotted Channel As A Gasdynamical Source Of Ignition And Flame Stabilization In The Supersonic Combustion Chamber

The original scheme of flame stabilization in the channel at close to cocurrent fuel supply for the fuel combustion at a high supersonic speed has been designed. Such solution provides high temperature of a stream in a zone of fuel-air mixture formation. Computational and experimental investigations of self-ignition and combustion of hydrogen were carried out in the model of combustor chamber with slotted channel (gasdynamical source of ignition) at Mach numbers 3.7 and 5.8 at the entrance. Tests have been performed in hot-shot wind tunnel IT-302M of ITAM SB RAS in a mode of the attached pipe. Numerical study has been performed on the basis of solving the full averaged Navier-Stokes equations, supplemented k-Q SST turbulence model. Configuration of the slotted channel has been designed with two variants of exit nozzle: with and without geometrical throat. It has been established that at the channel entrance two vortexes with high temperature have been appeared. Temperature has been keeping high in the channel with geometrical throat and at blocking of the slotted channel without throat. It was found that uniform subsonic stream in the channel with geometrical throat has been realized. The stream in the slotted channel without geometrical throat keeps supersonic but Mach number was lower than in the main channel. The structure of the flow at the slotted channel exit is significantly differs for this both cases.

Inertial forces affect fluid front displacement dynamics in a pore-throat network model.

The seemingly regular and continuous motion of fluid displacement fronts in porous media at the macroscopic scale is propelled by numerous (largely invisible) pore-scale abrupt interfacial jumps and pressure bursts. Fluid fronts in porous media are characterized by sharp phase discontinuities and by rapid pore-scale dynamics that underlie their motion; both attributes challenge standard continuum theories of these flow processes. Moreover, details of pore-scale dynamics affect front morphology and subsequent phase entrapment behind a front and thereby shape key macroscopic transport properties of the unsaturated zone. The study presents a pore-throat network model that focuses on quantifying interfacial dynamics and interactions along fluid displacement fronts. The porous medium is represented by a lattice of connected pore throats capable of detaining menisci and giving rise to fluid-fluid interfacial jumps (the study focuses on flow rate controlled drainage). For each meniscus along the displacement front we formulate a local inertial, capillary, viscous, and hydrostatic force balance that is then solved simultaneously for the entire front. The model enables systematic evaluation of the role of inertia and boundary conditions. Results show that while displacement patterns are affected by inertial forces mainly by invasion of throats with higher capillary resistance, phase entrapment (residual saturation) is largely unaffected by inertia, limiting inertial effects on hydrological properties behind a front. Interfacial jump velocities are often an order of magnitude larger than mean front velocity, are strongly dependent on geometrical throat dimensions, and become less predictable (more scattered) when inertia is considered. Model simulations of the distributions of capillary pressure fluctuations and waiting times between invasion events follow an exponential distribution and are in good agreement with experimental results. The modeling approach provides insights into the rich pore-scale dynamics of displacement fronts; these insights not only improve the basic understanding of these ubiquitous processes, but could shed light on solute dispersion and colloids mobilization at fronts and the mechanical consequences of passing fronts.

Aerodynamic Performance of Scramjet Inlet Models with a Single Strut

A series of three-dimensional sidewall compression-type scramjet inlet models with single struts were tested in a Mach 4 wind tunnel to investigate the influence of a strut between the sidewalls. With the sidewalls having a fixed shape, five struts of the same width but different lengths and wedge angles were mounted, so that the effect of each model could be compared with the same geometrical contraction ratio. The aerodynamic performances of inlets were determined based on the total pressure efficiency and the mass capture ratio at the geometrical throat. Also, flow visualizations were carried out to examine the flow structure, especially around the strut. Some of the results were compared with previously examined nonstrut models. The additional shock waves generated by the strut created a large separation on the sidewalls and the strut wall, resulting in the reduction of the total pressure efficiency and the capture ratio. Due to this separation, the shock wave was observed to be oscillatory. Measurement of fluctuating wall pressure was also carried out, which revealed that, in some regions, the pressure signals were asynchronous on the left and right sides of the strut. It seems that a small difference of the separation bubble structure on each sidewall initiated a cascade effect of pressure imbalance around the struts.

Downstream Ramjet-Mode Combustion in a Dual-Mode Scramjet Engine

Downstream combustion ramjet-mode operation in a dual-mode engine combustor was studied experimentally with a Mach 2.5 wind tunnel. In this downstream combustion ramjet mode, subsonic combustion was attained in the downstream straight-duct section without a geometrical throat. The experimental results are presented. The total temperature and the total pressure in the vitiation air heater were 800 K and 1.0 MPa, respectively. Pitot pressure and gas sampling were measured on the exit plane of the combustion model. The combustion conditions of the downstream combustion ramjet mode were compared with those of the usual upstream combustion ramjet mode. Better thrust performance and combustion were observed in the upstream combustion ramjet mode. However, in the case of the downstream combustion ramjet mode, injection of a larger amount of fuel was possible and a large impulse function was attained. The downstream straight duct was required for sufficient reaction. The fuel did not go far upstream in the separation region in pseudo-shock. Nomenclature ER = equivalence ratio M = Mach number Pi = wall pressure at the entrance of the combustor Pp = pitot pressure Pw = wall pressure x = distance from backward-facing step y = vertical distance from the bottom z = transverse distance from side wall surface ηc = combustion efficiency φ = equivalence ratio

Study about the flashing process through a metering expansion valve

In this paper, the results from an experimental campaign about the flashing process occurring through a metering manual expansion valve, are presented and discussed. The results involved the measurement at different lifts of the flow-rate flowing through the valve for three different refrigerants R22, R290 (propane) and R410A. First the experimental rig is described and the employed instrumentation is commented upon and the corresponding uncertainties analyzed. The test matrix is then presented and discussed. Then the geometry of the valve is described and an analysis about its geometric throat area as a function of the valve lift is also presented. The results of mass flow rate for different upstream pressure and temperature, subcooling, and downstream pressure are presented and the effect of each of those operating condition on the obtained mass flow rate is discussed. Different flow models are applied in order to evaluate the effective flow area of the valve at every valve lift and a comparison study is performed in order to elucidate which is the model which better explains the obtained results. Finally the results obtained with the best estimate model are shown for the full collection of tested points and the three tested refrigerants.

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