Total Pressure Probe Research Articles

Turbulent boundary layer separation control and loss evaluation of low profile vortex generators

The present paper analyses the results of a detailed experimental study on low profile vortex generators used to control the turbulent boundary layer separation on a large-scale flat plate with a prescribed adverse pressure gradient, typical of aggressive turbine intermediate ducts. This activity is part of a joint European research program on Aggressive Intermediate Duct Aerodynamics (AIDA). Laser Doppler Velocimetry and a Kiel total pressure probe have been employed to perform measurements in the test section symmetry plane and in cross-stream planes to investigate the turbulent boundary layer, with and without control device application. Velocity fields, Reynolds stresses, and total pressure distributions are analysed and compared for the controlled and non controlled flow conditions to characterize the mean flow behaviour. The detail and the accuracy of the measurements allow the evaluation of the deformation works of the mean motion in the test section symmetry plane. Normal and shear contributions of viscous and turbulent deformation works have been analysed and employed to explain the distribution of the total pressure loss. For the controlled flow the discussion of the flow field is extended considering the effects of the vortex generated in the cross-stream planes. The experimental data allow the evaluation of the global amount of losses, considering a balance of total pressure fluxes in the different measuring planes.

Numerical Investigation of the Influence of Real World Blade Profile Variations on the Aerodynamic Performance of Transonic Nozzle Guide Vanes

This paper addresses the issue of aerodynamic consequences of small variations in airfoil profile. A numerical comparison of flow field and cascade pressure losses for two representative repaired profiles and a reference new vane were made. Coordinates for the three airfoil profiles were obtained from the nozzle guide vanes of refurbished turboshaft engines using 3D optical scanning and digital modeling. The repaired profiles showed differences in geometry in comparison with the new vane, particularly near the leading and trailing edges. A numerical simulation was conducted using a commercial CFD code, which uses the finite volume approach for solving the governing equations. The computational predictions of the aerodynamic performance were compared with experimental results obtained from a cascade consisting of blades with the same airfoil profiles. The CFD analysis was performed for the cascade at subsonic inlet and transonic exit conditions. Boundary layer growth, wake formation, and shock boundary layer interactions were observed in the two-dimensional computations. The flow field showed the presence of shock waves downstream of the passage throat and near the trailing edges of the blades. A conspicuous change in flow pattern due to subtle variation in airfoil profile was observed. The calculated flow field was compared with the flow pattern visualized in the experimental test rig using the schlieren method. The total pressure calculation for the cascade exit showed an increase in pressure loss for one of the off-design profiles. The pressure loss calculations were also compared with the multihole total pressure probe measurement in the transonic cascade rig.

Unsteady Pressure Measurements With a Fast Response Cooled Probe in High Temperature Gas Turbine Environments

This paper presents the first experimental engine and test rig results obtained from a fast response cooled total pressure probe. The first objective of the probe design was to favor continuous immersion of the probe into the engine to obtain a time series of pressure with a high bandwidth and, therefore, statistically representative average fluctuations at the blade passing frequency. The probe is water cooled by a high pressure cooling system and uses a conventional piezoresistive pressure sensor, which yields, therefore, both time-averaged and time-resolved pressures. The initial design target was to gain the capability of performing measurements at the temperature conditions typically found at high pressure turbine exit (800–1100°C) with a bandwidth of at least 40 kHz and in the long term at combustor exit (2000 K or higher). The probe was first traversed at the turbine exit of a Rolls-Royce Viper turbojet engine at exhaust temperatures around 750°C and absolute pressure of 2.1 bars. The probe was able to resolve the high blade passing frequency (≈23 kHz) and several harmonics of up to 100 kHz. Besides the average total pressure distributions rom the radial traverses, phase-locked averages and random unsteadiness are presented. The probe was also used in a virtual three-hole mode yielding unsteady yaw angle, static pressure, and Mach number. The same probe was used for measurements in a Rolls-Royce intermediate pressure burner rig. Traverses were performed inside the flame tube of a kerosene burner at temperatures above 1600°C. The probe successfully measured the total pressure distribution in the flame tube and typical frequencies of combustion instabilities were identified during rumble conditions. The cooling performance of the probe is compared with estimations at the design stage and found to be in good agreement. The frequency response of the probe is compared with cold shock-tube results and a significant increase in the natural frequency of the line-cavity system formed by the conduction cooled screen in front of the miniature pressure sensor were observed.

원심 다익홴의 유동에 대한 컷 오프 각도의 영향

This study investigated on details of flow characteristics of a multi-blade fan for domestic ventilation. Experiments and analysis were carried out to describe on flow pattern with variations of cut-off angle near the scroll housing throat, which were performed by PIV measurement for the flow field and by total pressure probes. The stagnation point at cut-off region of the fan moves to the exit of the scroll housing as the cut-off angle increases. The movement of stagnation point and the variation of throat area of the scroll housing influence to the distribution of velocity magnitude at the exit of the fan. Furthermore, a large distortion of the velocity distribution at the scroll exit causes to increase mixing loss along the flow path.

Simultaneous Measurement of Velocity and Fluctuating Pressure in a Turbulent Wing-tip Vortex Using Triple Hot-film Sensor and Miniature Total Pressure Probe

We have developed a probe-system for simultaneous measurement of three velocity components and pressure in turbulent flows. A miniature total pressure probe is placed adjacent to the sensors of a triple hot-film probe in order to achieve the spatial resolution which is equivalent to that of the triple hot-film probe itself. The instantaneous static pressure is calculated from measured velocity and total pressure by means of a newly developed processing method based on the Bernoulli equation for unsteady flows. The measurements were undertaken in a turbulent wing-tip vortex flow. The look-up table method is employed for the calibration of the hot-film probe so accurate velocity data could be obtained over a wide range of the flow-attack angles. It is also demonstrated that the present probe-system is capable of measuring fluctuations in both velocity and pressure in the 20–650 Hz frequency range. The distribution of the fluctuating pressure obtained by this indirect method is in good agreement with the results from direct measurements of static pressure, demonstrating the promising performance of the present method. Furthermore, an improvement in the ability to make measurements of the velocity–pressure correlation across the wing-tip vortex is achieved. This improvement is possible because the effects of lateral velocity components are properly taken into account in the present formulation. The investigation regarding the transport equation budget for turbulent kinetic energy shows an anomalous structure of turbulence in this flow, mainly due to the meandering of the vortex, and the measurement of pressure diffusion is found to play an important role in the characterization of this kind of flow.

Fast-response total pressure probe for turbomachinery application

To evaluate the accurate performance and characteristics of turbomachinery, it is important to measure the unsteady flow phenomena downstream of the rotating blades. This paper presents the development of a fast-response total pressure probe for the measurement of the total pressure field at the exit of blades. The result of measurement in a one-stage axial turbine is also presented. The fast-response total pressure probe is fabricated by installing a fast-response pressure sensor in the cylindrical head of the probe. In terms of simplicity of the measurement system and data reduction method, this method is more competitive over established methods that use more than four sensors. The probe is applied to the one-stage axial turbine in order to measure the instantaneous total pressure downstream of rotor blades. The measured instantaneous signal is decomposed to obtain the blade-to-blade pressure distribution. The pressure distribution due to blade passing is clearly captured. Due to the loss generation in the casing region, the total pressure and its amplitude of fluctuation by the blade passing are lower in the shroud and hub region than in mid-span. The total pressure distribution at the exit of the rotor blade is found to be slightly different from blade to blade due to the geometric difference and the different relative positions of the rotor blades and stator vanes. The developed probe successfully measures the accurate total pressure distribution at the rotor exit, and allows the evaluation of the loss distribution and the accurate performance of turbomachinery.

Behavior of shock trains in a diverging duct

A shock train inside a diverging duct is analyzed at different pressure levels and Mach numbers. Nonreactive pressurized cold gas is used as fluid. The structure and pressure recovery inside the shock train is analyzed by means of wall pressure measurements, Schlieren images and total pressure probes. During the course of the experiments, the total pressure of the flow, the back pressure level and the Mach number upstream of the compression region have been varied. It is shown that the Reynolds number has some small effect on the shock position and length of the shock train. However, more dominant is the effect of the confinement level and Mach number. The results are compared with analytical and empirical models from the literature. It was found that the empirical pseudo-shock model from Billig and the analytical mass averaging model from Matsuo are suitable to compute the pressure gradient along the shock train and total pressure loss, respectively.

Experimental Investigation of Boundary Layer Characteristics on Blade Surface under Different Inlet Flow Conditions

In this paper, an experimental study is conducted on cascade boundary layer under different inlet conditions. New method is used to measure the total pressure in blade surface boundary layer directly using total pressure probe. Total pressure in both suction and pressure surfaces are acquired at different inlet conditions by changing incidence angle and inlet Mach number. In addition, a series of parameters related to boundary layer characteristics are calculated. The objective of the experiment is to investigate the influence of inlet flow conditions on them. The results indicate that influence of incidence angle is significant when other conditions are the same. Displacement thickness, momentum thickness as well as other parameters display some disciplines for variation. In contrast, inlet Mach number has only a small influence in that boundary layer becomes a litter thinner with increasing Mach number. Comparisons of experimental results with theoretical expectations demonstrate that the method in this experiment is effective and reliable.

Some Considerations about the Rotating Cell Structure

Rotating stall instability in axial flow compressors arises when the mass flow is reduced at constant rotational speed. Despite the number of experimental and theoretical work already published in the scientific literature, many questions still remain unanswered. A complete model that could be of help both in the design process and in the modelling process of existing engines is not yet available. A fundamental research program has been carried out at the Department of Mechanical Engineering of the University of Cagliari with the scope of reaching a better understanding of the basic flow field structures and of the global performance in stall. The cinematic structure of the cell during abrupt full 1‐cell stall in a two‐stage axial flow compressor with IGV has been analysed with the aid of hot‐wire anemometer and of a total pressure probe. The results have revealed interesting features about the cell flow structures and their variations along the stalled performance branch of the compressor. The present paper aims at pursuing further the flow analysis including the investigation of upstream flow field. Therefore, the inlet duct has been extended to perform the measurements. The complete flow field measurements will allow to obtain a complete cinematic description of the cell.

A Conceptual Design Study for a New High Temperature Fast Response Cooled Total Pressure Probe

The present paper proposes a concept for a water-cooled high temperature unsteady total pressure probe intended for measurements in the hot sections of industrial gas turbines or aero-engines. This concept is based on the use of a conventional miniature piezoresistive pressure sensor, which is located at the probe tip to achieve a bandwidth of at least 40 kHz. Due to extremely harsh conditions and the intention to immerse the probe continuously into the hot gas stream, the probe and sensor must be heavily cooled. The short term objective of this design is to gain the capability of performing measurements at the temperature conditions typically found at high pressure turbine exit (1100–1400 K) and in the long term at combustor exit (2000 K or higher).

Time Resolved Experimental Investigations of an Axial Compressor With Casing Treatment

A casing treatment with axial and radial skewed slots ending in a plenum chamber has experimentally been investigated at a highly subsonic axial compressor stage. The aim was to investigate the physical phenomenon of this treatment family that is responsible for the stabilization of the blade passage flow and the drop in efficiency mostly observed. The experimentally gained performance results of this configuration showed an extension of the operating range by approximately 50%, while the efficiency for design conditions is reduced by 1.4%. Apart from this, operating points at part load conditions have been observed nearly without any loss in efficiency. The detailed flow analysis is performed by means of results from a 3D pneumatic probe with temperature sensor and a dynamic total pressure probe. The focus of the investigations is on the incidence flow to the compressor rotor, the tip clearance vortex flow in combination with the wall stall separation region and the blade stall due to suction side separation. The casing treatment configuration is investigated with a special interest in detecting those effects which have an impact on the stability and the compressor overall efficiency, including the interaction of the rotor and the stator flow fields.

Effects of blade sweep on the performance characteristics of axial flow turbomachinery rotors

Experimental studies were carried out in order to survey the performance and efficiency aspects of spanwise constant forward and backward sweep in axial flow rotors of low aspect ratio (AR) blading for incompressible flow, for part-load, near-design, and overload operational ranges. The experiments involved overall performance tests, laser Doppler velocimetry, and stationary total pressure probe measurements. The overall performance data and pitchwise averaged and resolved flow characteristics were evaluated in detail. For moderate and high flow rates, it was pointed out that positive or negative sweep tends to reduce or increase the blade load in the vicinity of the endwalls, respectively. It has been concluded that the loss-modifying effect of sweep can be judged by considering the three-dimensional viscous phenomena, and the influence of sweep on local blade efficiency depends on the balance of changes in blade load and losses. For low flow rates, forward sweep was found beneficial over the entire span from the aspect of improved stall margin and efficiency. The influence of AR on the performance reducing effect of sweep was studied on the basis of literature data.

Characteristics of the AT-303 hypersonic wind tunnel. Part 1. Velocity Fields

The structure and principle of operation of a new wind tunnel AT-303 with adiabatic compression are described. Results of systematic investigations are presented in terms of velocity distributions both at the nozzle exit and in the region where the models are located. The velocity fields are obtained with the use of total pressure probes in the ranges of Mach numbers from 7.6 to 19.7 and Reynolds numbers per meter Re1 = (0.25−3.64)·107.

Heat Transfer, Pressure Loss and Flow Field Measurements Downstream of Staggered Two-Row Circular and Elliptical Pin Fin Arrays

This paper presents the results of heat transfer, total pressure loss, and wake flow field measurements downstream of two-row staggered elliptical and circular pin fin arrays. Two different types of elliptical fins are tested, i.e., a Standard Elliptical Fin (SEF) and a fin that is based on NACA four digit symmetrical airfoil shapes (N fin). The results are compared to those of a corresponding circular pin fin array. The minor axis lengths for both types of elliptical fins are kept equal to the diameter of the circular fins. Experiments are performed using Liquid Crystal Thermography and total pressure probe wake surveys in a Reynolds number range of 18 000 and 86 000 as well as Particle Image Velocimetry (PIV) measurements at ReD=18 000. The pin fins had a height-to-diameter ratio of 1.5. The streamwise and the transverse spacings were equal to one circular fin diameter, i.e., S/D=X/D=2. For the circular fin array, average Nusselt numbers on the endwall within the wake are about 27% higher than those of SEF and N fin arrays. Different local heat transfer enhancement patterns are observed for elliptical and circular fins. In terms of total pressure loss, there is a substantial reduction in case of SEF and N fins. The loss levels for the circular fin are 46.5% and 59.5% higher on average than those of the SEF and N fins, respectively. An examination of the Reynolds analogy performance parameter show that the performance indices of the SEF and the N fins are 1.49 and 2.0 times higher on average than that of circular fins, respectively. The thermal performance indices show a collapse of the data, and the differences are much less evident. Nevertheless, N fins still show slightly higher thermal performance values. The wake flow field measurements show that the circular fin array creates a relatively large low momentum wake zone compared to the SEF and N fin arrays. The wake trajectories of the first row of fins in circular, SEF and N fin arrays are also different from each other. The turbulent kinetic energy levels within the wake of the circular fin array are higher than those for the SEF and the N fin arrays. The transverse variations in turbulence levels correlate well with the corresponding local heat transfer enhancement variations.

Unsteady Pressure Measurements around Rotor of an Axial Flow Fan under Stable and Unstable Operating Conditions

This study presents some measurement results on the unsteady pressure fields around rotor under stable and unstable operating conditions of an axial flow fan. The unsteady static pressure of rotor passage was measured by using high frequency pressure transducers mounted on the casing wall. The measurements on the unsteady total pressure at rotor inlet and outlet were also conducted with specially designed high frequency total pressure probe. Double Phase-Locked Ensemble Averaging Technique was used for analysis of pressure fluctuations around the rotor at rotating stall onset point. From the results, the unsteady pressure fields during stable and unstable operations of the axial fan were investigated and compared with each other. Particularly one period of rotating stall could be divided into two regions, stalled flow and unstalled flow region respectively. Furthermore the former could be also classified into two zones, bubbled and disturbed region by their features. The flow characteristics for each zone were described in detail and the static and total pressure fields were also analyzed in terms of the pressure distribution along pressure side and suction side on the blade tip profile.

Experimental investigations of rotating flow instabilities in the last stage of a low-pressure model steam turbine during windage

Experimental investigations on two different three-stage low-pressure (LP) model turbines operated on steam during windage are presented. At very low flows, strain gauges identified that the two turbines exhibit different dynamic stress characteristics of the last-stage rotor blades. The first turbine experienced the highest vibrations from aerodynamic excitations at the resonant frequency of the second mode in a narrow operating range around 13 per cent of the design flow, whereas the second turbine experienced no resonance up to the fourth mode and has the highest dynamic loading at zero flow. Specially developed high-response total pressure probes have been used to determine the cause of the aerodynamic excitation from traverses of the unsteady pressure field at three planes in the last stage. The experimental data for both turbines show that unsteady pressure disturbances steadily grow when the flow is reduced below 25 per cent of design flow and that the excitations are strongest in the axial gap between the guide vane and the rotor of the last stage near the outer casing. Detailed analysis shows that high-amplitude disturbances occur at distinct frequencies and that these rotate in the circumferential direction at a fraction of the rotor speed. Comparison of the pressure signals measured at two circumferential locations on the casing of the second turbine confirmed the characteristic frequency pattern to be a so called ‘rotating instability’. This unsteady phenomenon arising from the tip leakage flow has previously been observed in axial flow fans and compressors and is demonstrated for the first time here in a turbine operating at very low flow rates.

Transonic and Low Supersonic Flow Losses of Two Steam Turbine Blades at Large Incidences

A linear cascade experiment was conducted to investigate transonic and low supersonic flow losses of two nozzle blades for the steam turbines. In the experiment, flow incidences were changed from −34° to 35° and exit Mach numbers were varied from 0.60 to 1.15. Tests were conducted at Reynolds numbers between 7.4×105 and 1.6×106. Flow visualization techniques, such as shadowgraph, Schlieren, and surface color oil were used to document the flows. Measurements were made by using downstream traverses with Pitot probe, upstream total pressure probe, and sidewall static pressure taps. The losses were found to be rather constant at subsonic flows. At transonic and low supersonic flows, the losses increased steeply. The maximum relative increase of the losses was near 700% when the Mach numbers increased from 0.6 to 1.15. However, the maximum relative increase of the losses was only about 100% due to very large variation of incidences. It is important to note that the effect of Mach numbers on losses was much greater than that due to the very large incidences for the transonic and low supersonic flows. A frequently used loss correlation in the literature is found not suitable to predict the losses of the tested blades for the transonic and low supersonic flows. From the current experimental data and some data in the literature, a new correlation for the shock related losses is proposed for transonic and low supersonic flows of turbine cascades. Comparison is made among the existing correlation and the new correlation, as well as the data of the current two cascades and other three turbine cascades in the literature. Improved agreement with the experimental data of the five cascades is obtained by using the new correlation as compared with the prediction by using the frequently used loss correlation in the literature.

Atmospheric Chemistry in Giant Planets, Brown Dwarfs, and Low-Mass Dwarf Stars: I. Carbon, Nitrogen, and Oxygen

The chemical species containing carbon, nitrogen, and oxygen in atmospheres of giant planets, brown dwarfs (T and L dwarfs), and low-mass stars (M dwarfs) are identified as part of a comprehensive set of thermochemical equilibrium and kinetic calculations for all elements. The calculations cover a wide temperature and pressure range in the upper portions of giant planetary and T-, L-, and M-dwarf atmospheres. Emphasis is placed on the major gases CH 4, CO, NH 3, N 2, and H 2O but other less abundant gases are included. The results presented are independent of particular model atmospheres, and can be used to constrain model atmosphere temperatures and pressures from observations of different gases. The influence of metallicity on the speciation of these key elements under pressure–temperature (P–T) conditions relevant to low-mass object atmospheres is discussed. The results of the thermochemical equilibrium computations indicate that several compounds may be useful to establish temperature or pressure scales for giant planet, brown dwarf, or dwarf star atmospheres. We find that ethane and methanol abundance are useful temperature probes in giant planets and methane dwarfs such as Gl 229B, and that CO 2 can serve as a temperature probe in more massive objects. Imidogen (NH) abundances are a unique pressure-independent temperature probe for all objects. Total pressure probes for warmer brown dwarfs and M dwarfs are HCN, HCNO, and CH 2O. No temperature-independent probes for the total pressure in giant planets or T-dwarf atmospheres are identified among the more abundant C, N, and O bearing gases investigated here.

広がりノズルからの自励振動とジェット騒音との関連性

In this study, the relation between vibration the self-induced oscillation and noise of jet flow issuing from divergent nozzle were studied by the flow visualization, pressure measurement and noise measurement. The instantaneous flow visualization ware taken by schlieren method. The pressure fluctuation of total pressure was measured by a pressure transducer assembled into a total pressure probe and further, the noise of sound pressure level for the linear scale was measured. The frequency of the pressure and noise was analyzed by FFT analyzer. As a result, it was found that the jet noise was influenced by the total pressure fluctuation.

Effects of Gurney Flaps on a NACA0012 Airfoil

Experimental measurements of surface pressure distributions and wake profiles were obtained for a NACA0012 airfoil to determine the lift, drag, and pitching-moment coefficients for various configurations. The addition of a Gurney flap increased the maximum lift coefficient from 1.37 to 1.74, however there was a drag increment at low-to-moderate lift coefficient. In addition, the boundary layer profile measurements were taken using a rake of total pressure probes at the 90% chord location on the suction side. The effective Gurney flap height is about 2% of chord length, which provides the highest lift-to-drag ratio among the investigated configurations when compared with the clean NACA0012 airfoil. In this case, the device remains within the boundary layer.