Fast-response Pressure Transducers Research Articles

Effect of Stage Loading and Tip Clearance on Tip Clearance Flows in a Centrifugal Compressor

The present paper reports measurement of periodic static pressures on the casing over the impeller of a low speed centrifugal compressor at different stage loadings and tip clearances. The periodic pressure data is acquired employing the technique called PLEAT. Experiments are carried out in a low speed centrifugal compressor with unshrouded impeller. Twelve holes are provided on the casing in which adapters are attached. A fast response pressure transducer, Kulite XCS-062 5pisd is inserted in the adapters. The pressure transducer measures pressure in terms of voltage. The voltage output is very low. The output voltage is amplified and converted to digital values for further processing of data. Sample results are presented and interpreted. From the measured static pressure, tip clearance flows can be easily identified. Effect of tip clearance and stage loading on tip clearance flows is highlighted.

Investigations of shock–boundary layer interaction dynamics using high-bandwidth pressure field imaging

The large-scale pulsations of shock-induced separation with length scale that significantly exceeds the incoming boundary layer thickness are investigated. The shock–boundary layer interaction (SBLI) unit is generated by an inward-turning axisymmetric compression ramp at an inflow Mach number of 2.5. A substantial region surrounding the centre azimuth exhibited mean and dynamic flow features that are consistent with two-dimensional separation. Two-dimensional highly resolved maps of surface pressure field are obtained using fast-response pressure-sensitive paint fluorescence imaging at 40 kHz repetition rate. The measurement domain covered significant regions of the incoming boundary layer through the relaxing boundary layer downstream of the reattachment as well as over 25 boundary layer thicknesses in the azimuthal direction. These measurements provide new insights into the spanwise coupling of the SBLI unit in addition to its inherent dynamics. The power spectral density (PSD) of the centreline pressure exhibits very good agreement with theoretical models and complementary measurements using fast-response pressure transducers, which served to validate the pressure field measurements. Detailed examination of the PSD reveals strong agreement with the literature, which includes the peak Strouhal number of the separation and reattachment shock motions as well as the downward frequency shift along the separation bubble. Furthermore, the pressure fluctuation maps reveal streamwise-elongated structures just downstream of the ramp leading edge that persist well downstream of the reattachment. A time sequence of conditional average pressure fluctuation maps is constructed surrounding isolated pressure excursions in the intermittent region. This sequence, along with two-point cross correlation analysis, provides critical information about the flow processes that drive the separation bubble pulsations in the SBLI units with large separation scales. Overall, the imbalance in the mass within the separation bubble appears to be the critical mechanism that drives the separation bubble pulsations. Furthermore, the pressure perturbations originating at azimuthally offset locations are also observed to influence the separation bubble dynamics.

Dynamic Stall Investigation on a Rotating Semielastic Double-swept Rotor Blade at the Rotor Test Facility Göttingen

Experimental investigations of three-dimensional dynamic stall on a four-bladed Mach-scaled semielastic rotor with an innovative double-swept rotor blade planform are presented. The study focuses on the coupling between the aeroelastic behavior of the blade and the underlying aerodynamics. Blade bending moment and flap displacement measurements were conducted using strain gauges and optical tracking of blade tip markers. The aerodynamic behavior was characterized by means of unsteady surface pressure measurements using unsteady pressure-sensitive paint (iPSP) across the outer 65% of the blade span and fast response pressure transducers at discrete locations. Different cyclic-pitch settings were investigated at a rotation frequency of frotor = 23.6 Hz that corresponds to blade tip Mach and Reynolds numbers of Mtip = 0.282– 0.285 and Retip = 5.84-5.95 ×10 5. The findings reveal a detailed insight into the nonlinear behavior in the flap movement during downstroke. iPSP and pressure transducer data indicate that this nonlinear flap behavior is caused by a radially phase-shifted dynamic stall process at the forward and backward swept part of the blade.

Experimental study of hypersonic traveling crossflow instability over a yawed cone

Boundary layer transition is an important factor in the design process of hypersonic vehicles, and crossflow instability is a critical mode. In this paper, the traveling crossflow instability in a laminar boundary layer at Mach 6 is experimentally investigated. Experiments were performed over a 7∘ half-angle cone at 6∘ angle of attack, with the unit Reynolds number of Re∞=5.75×106 m−1. Flush-mounted fast-response pressure transducers and time-averaging temperature-sensitive paint were used to capture the traveling crossflow instability and the stationary crossflow instability, respectively. The traveling crossflow waves with characteristic frequencies between 10 and 25 kHz are detected in the planes with the azimuthal angle between 23∘ and 143∘ away from the leeward ray. The three-dimensional amplitude growth was obtained. In general, the amplitude growth of the traveling crossflow instability is roughly leeside-forward and windside-aft. Although the traveling crossflow instability grows earlier on the leeward side, it has a larger amplitude on the windward side. Compared with the stationary mode, the traveling mode grows faster and reaches the highest amplitude earlier. In addition, a high-frequency instability with characteristic frequencies between 100 and 200 kHz is observed. It may be the secondary instability of the traveling crossflow waves, but whether it is true or not needs further experimental and numerical studies.

Dynamic-stall measurements using time-resolved pressure-sensitive paint on double-swept rotor blades

The study presents an optimized pressure-sensitive paint (PSP) measurement system that was applied to investigate unsteady surface pressures on recently developed double-swept rotor blades in the rotor test facility at the German Aerospace Center (DLR) in Göttingen. The measurement system featured an improved version of a double-shutter camera that was designed to reduce image blur in PSP measurements on fast rotating blades. It also comprised DLR’s PSP sensor, developed to capture transient flow phenomena (iPSP). Unsteady surface pressures were acquired across the outer 65% of the rotor blade with iPSP and at several radial blade sections by fast-response pressure transducers at blade-tip Mach and Reynolds numbers of mathrm {M}_mathrm{tip} = 0.282-0.285 and mathrm {Re}_mathrm{tip}= 5.84-5.95 times 10^5. The unique experimental setup allowed for scanning surface pressures across the entire pitch cycle at a phase resolution of {0.225},{mathrm{deg}} azimuth for different collective and cyclic-pitch settings. Experimental results of both investigated cyclic-pitch settings are compared in detail to a delayed detached eddy simulation using the flow solver FLOWer and to flow visualizations from unsteady Reynolds-averaged Navier–Stokes (URANS) computations with DLR’s TAU code. The findings reveal a detailed and yet unseen insight into the pressure footprint of double-swept rotor blades undergoing dynamic stall and allow for deducing “stall maps”, where confined areas of stalled flow on the blade are identifiable as a function of the pitch phase.Graphical abstract

Experimental visualization and characteristics of bubble nucleation during rapid decompression of supercritical and subcooled carbon dioxide

The rapid decompression of sub-cooled and supercritical carbon dioxide was investigated experimentally using schlieren and regular high-speed photography. An optically accessible expansion tube was designed and constructed. A test campaign with initial pressures ranging from 6.45–9.02 MPa and initial temperatures from 297–308 K was performed. High-speed schlieren imagery was used to visualize the decompression waves, evaporation waves, and an oscillatory phase change region present in the blowdown process. The time history of pressures at different locations along the expansion tube were recorded using fast response pressure transducers. Image and pressure signal analyses were performed to extract decompression wave speeds and evaporation wave speeds. Results were compared to analytical models and other experiments reported in the literature. The relationship between pressure undershoot and wave speed to the initial conditions was analysed.

Prediction of Transient Pressure Fluctuations within a Low-Pressure Turbine Cascade Using a Lanczos-Filtered Harmonic Balance Method

Unsteady pressure fluctuations measured by fast-response pressure transducers mounted in a low-pressure turbine cascade are compared to unsteady simulation results. Three differing simulation approaches are considered, one time-integration method and two harmonic balance methods either resolving or averaging the time-dependent components within the turbulence model. The observations are used to evaluate the capability of the harmonic balance solver to predict the transient pressure fluctuations acting on the investigated stator surface. Wakes of an upstream rotor are generated by moving cylindrical bars at a prescribed rotational speed that refers to a frequency of f∼500 Hz. The excitation at the rear part of the suction side is essentially driven by the presence of a separation bubble and is therefore highly dependent on the unsteady behavior of turbulence. In order to increase the stability of the investigated harmonic balance solver, a developed Lanczos-type filter method is applied if the turbulence model is considered in an unsteady fashion.

Experimental investigation of a free surface oscillation in a model pump-turbine

Hydro power plants provide balancing services to the power grid and the amount is expected to increase in the near future. One of these services is the compensation of reactive power and the regulation of the power factor. Therefore, hydro power units equipped with synchronous machines are operated in so called synchronous condenser mode. During this operation the runner is dewatered and rotates in air above a free surface at the draft tube cone. As a result, a free surface oscillation is excited under some conditions. This paper deals with the experimental investigation of this flow phenomenon. Therefore, model tests are conducted at the hydraulic laboratory of the Institute of Energy Systems and Thermodynamics at TU Wien. The investigated model pump-turbine is equipped with a draft tube cone made of acrylic glass for optical investigations. In addition to the standard instrumentation, fast response pressure transducers are installed at different planes and positions. The machine setup and measurement conditions aim at the investigation of the free surface oscillation and shall prevent other influences. The experiments are performed for both runner rotation directions, different initial water levels at the draft tube and different densimetric Froude numbers. High speed camera recordings are conducted and the amplitude of the water formation is determined by post processing the captured images. In periodic state the frequency of the free surface oscillation is extracted out of the pressure signals. Finally general trends and dependencies are deduced from the acquired data.

On High-Resolution Pressure Amplitude and Phase Measurements Comparing Fast-Response Pressure Transducers and Unsteady Pressure-Sensitive Paint

AbstractThis paper presents the simultaneous application of fast-response pressure transducers and unsteady pressure-sensitive paint (unsteady PSP) for the precise determination of pressure amplitudes and phases up to 3000 Hz. These experiments have been carried out on a low-pressure turbine blade cascade under engine-relevant conditions (Re, Ma, and Tu) in the high-speed cascade wind tunnel. Periodic blade/vane interactions were simulated at the inlet to the cascade using a wake generator operating at a constant perturbation frequency of 500 Hz. The main goal of this paper is the detailed comparison of amplitude and phase distributions between both flow sensing techniques at least up to the second harmonic of the wake generator’s fundamental perturbation frequency (i.e., 1000 Hz). Therefore, a careful assessment of the key drivers for relative deviations between measurement results as well as a detailed discussion of the data processing is presented for both measurement techniques. This discussion outlines the mandatory steps which were essential to achieve the quality as presented down to pressure amplitudes of several Pascal even under challenging experimental conditions. Apart from the remarkable consistency of the results, this paper reveals the potential of (unsteady) PSP as a future key flow sensing technique in turbomachinery research, especially for cascade testing. The results demonstrate that PSP was able to successfully sense pressure dynamics with very low fluctuation amplitudes down to 8 Pa.

Experimental study of crossflow instability in a Mach 6 delta wing flow

In this paper, the traveling crossflow instability in the boundary layer on the windward side of a delta wing is studied. The experiments are carried out in a Mach 6 low-noise wind tunnel, with the angles of attack of the model being 5° and 10°, and the Reynolds number being in a range of 2.43 × 10<sup>6</sup>–14.21 × 10<sup>6</sup> m<sup>–1</sup>. The wall fluctuation pressure is measured by fast-response Kulite pressure transducers. The power spectrum density (PSD) analysis is conducted to obtain the disturbance waves' development process in the boundary layer. The temperature-sensitive paints (TSPs) and nano-tracer based planar laser scattering (NPLS) technique are also used. From the TSP results, the boundary layer transition near the leading edge of the delta wing is smooth and parallel to the leading edge. A peak around 10 kHz in power spectrum density is detected by the fast-response pressure sensor, which may be caused by the traveling crossflow waves. To verify this dominant mode, an NPLS image in the plane of <i>n</i> = 36 mm is obtained. The shapes of vortex structures correspond to the shapes of the crossflow vortices from the numerical simulation. Only when the boundary layer is laminar can the traveling crossflow wave signal be observed from the PSD curves. When the boundary layer is at a transitional or turbulent phase, the low-frequency component is dominant in the PSD curve. With the increase of Reynolds number, the characteristic frequency of the crossflow wave increases, and the wave’s amplitude first increases and then decreases. Moreover, the angle of attack effect is obtained. The increasing of the angle of attack can make the traveling crossflow wave grow faster and saturate, attenuate at the position closer to the leading edge of the delta wing or at a lower Reynolds number. By sensor pairs composed of three Kulite transducers, the phase velocity and the propagation angle of the traveling crossflow wave are investigated. The dimensionless phase velocities of the traveling wave are in ranges of 0.24–0.26 and 0.26–0.32 at 5° and 10° angles of attack, respectively. The propagation angles are at 50°–60° and 40°–55° at the angles of attack of 5° and 10°, respectively. At a larger angle of attack, the traveling wave’s propagation angel is smaller, but the phase velocity is bigger. It may be because the spanwise pressure gradient is higher at a larger angle of attack, and then the crossflow velocity is stronger.

Experimental investigation of nose-tip bluntness effects on the hypersonic crossflow instability over a cone

The nose-tip bluntness effects on the hypersonic crossflow instability were experimentally investigated over a 7∘ half-angle straight cone at 6∘ angle of attack. The experiments were conducted in a Mach 6 wind tunnel at Reynolds number of 7.94×106 m−1 using temperature-sensitive paint (TSP) and fast-response pressure transducers. The wavelet transform was applied to obtain the most amplified wavenumber of the stationary crossflow vortices. The stationary crossflow vortices with wavenumber of 40–60 and the traveling crossflow waves with characteristic frequencies of 9–20 kHz have been detected. TSP results indicate that the bluntness has little influence on the boundary layer transition location with a small nose-tip radius (smaller than 1.5 mm), while the transition front moves backward obviously with a larger nose-tip radius (larger than 2.5 mm). The developments of two types of crossflow instability, the stationary mode and the traveling mode, are both not sensitive to the bluntness when the nose-tip radius is below 1.5 mm. And the growth and saturation of the two types crossflow mode are both inhibited under larger bluntness. Besides, possible explanations for how the bluntness affects the transition location and crossflow instability is discussed.

Hypersonic Fluid–Structure Interactions in Compression Corner Shock-Wave/Boundary-Layer Interaction

The fluid–structure interaction of a flexible panel exposed to a ramp-induced shock-wave/boundary-layer interaction (SWBLI) at Mach 6 is investigated experimentally for transitional and turbulent incoming boundary layers. Panel deformations are measured using photogrammetry enabled by a new marker-tracking routine, whereas pressure fluctuations are obtained with fast-response piezoresistive pressure transducers. The significance of aerothermal heating is evident in the nonlinear panel response: enhanced static deformations and frequency shifting are consistent with a temperature differential between the panel and its support structure, which induces compressive thermal strain and flexural softening. Time-domain and modal vibratory behavior are correlated to the SWBLI environment, and shear-layer reattachment near antinodes of certain mode shapes is identified as a source of enhanced panel excitation. Comparison with companion rigid-ramp experiments shows evidence of feedback into the downstream flowfield regime.

Unsteady Flow Features Across Different Shock/Boundary-Layer Interaction Configurations

An experimental study has been conducted to investigate unsteady flow phenomena observed within various two-dimensional configurations of shock/boundary-layer interactions. Six configurations have been tested in Mach 2 flow: and 20° compression ramps, and incident shock reflections from , 8, 9, and 10° shock generators; Reynolds numbers in each case are . The flow is assessed using an array of fast-response pressure transducers in conjunction with a high-repetition rate particle image velocimetry system. Development of the mean flow structures early in each interaction is observed to be consistent with the free interaction concept. Unsteady wall-pressure energy content at frequencies above those associated with the characteristic low-frequency shock motion also shows significant similarities in the vicinity of the shock foot. Results confirm that this low-frequency peak is not associated with a narrow-band forcing mechanism from either upstream or downstream, but rather a characteristic frequency that varies with interaction strength, which describes the flow’s dynamic response. These findings support various models published in the literature that have sought to explain the source of low-frequency unsteady shock motion.

Unsteady measurement of core penetration flow caused by rotating geometric non-axisymmetry in a turbine rotor-stator disc cavity

The existence and causes of the deep ingress into the core region of a turbine rotor-stator disc cavity, or core penetration flow, generated by a rotating non-axisymmetric geometry have been investigated experimentally. In a low-speed, low-expansion ratio, single-stage, cold turbine test facility, time-resolved tangential and radial velocities in the cavity have been measured with 2-D hot-wire anemometers. In addition, time-resolved static pressures on the stator disc have been measured with fast response pressure transducers, and unsteady cavity velocity field in the absolute frame has been measured using Particle Image Velocimetry (PIV). Rotating geometric non-axisymmetry leads to an unsteady radial pressure gradient in the disc cavity. A time lag in the tangential velocity adjustment to the variation in the radial pressure gradient results in a net radial force, causing core penetration flow. A first-harmonic rotating geometric non-axisymmetry makes the core penetration flow to occur three times per revolution (twice when the cavity exit pressure increases and once when the cavity exit pressure decreases) and to revolve at the disc’s rotational speed. Core penetration flow due to the rotating geometric asymmetry is not affected by variations in the annulus flow coefficient or rotational Reynolds number but is weakened by increasing purge air flow rate.

Flame Propagation in a Narrow Closed Channel: Effects of Aspect Ratios, Blockage Ratio, and Mixture Reactivity on Flame Speed and Pressure Dynamics

ABSTRACT A set of experiments is conducted to study flame acceleration in a narrow rectangular cross-section channel in order to examine the effects of mixture reactivity, channel height, and blockage ratio (BR). The channel had an ignition chamber at one end and a dump chamber at the opposite end. The initial conditions are mixtures of hydrogen-air at 1 bar and 293 K. Five different compositions of 10, 20, 30, 40, and 50% hydrogen volume fraction in air are studied. The rectangular chamber is 500 mm long and 40 mm in width, with an adjustable height from 2 to 10 mm. To study the relation between channel height and BR, combustion behaviors with three different B.R. (BR of 0.1, 0.5, and 0.9), 5 different obstacle sizes, and different channel heights (from 2–10 mm) are presented. Pressure waves are monitored with seven fast response pressure transducers placed along the chamber, and the flame propagation is observed with a high-speed Schlieren imaging system. There was no detonation observed during this series of experiments. The results show that the flame propagation is affected by both the channel dimensions and the obstacles, with a maximum flame velocity achieved for partially obstructed channels. The evolution of the flame’s propagation and the pressure dynamics through the various obstructed and unobstructed channels, including the dynamics contributed by the ignition and dump chamber, are discussed.

Roles and mechanisms of casing treatment on different scales of flow instability in high pressure ratio centrifugal compressors

The development of single-stage high pressure ratio centrifugal compressors for small gas turbine engines is hindered by compressor flow instabilities with various temporal and spatial scales. In this paper, the effect of a self-recirculation casing treatment device for suppressing various scales of flow instability has been discussed. Rig tests of a high pressure ratio centrifugal compressor with and without a casing treatment device are performed, and transient pressure signals are measured at endwalls by fast response pressure transducers. By analyzing measured pressure signals in both time and frequency domain, the casing treatment device is found effective in eliminating rotating instability and stall at the impeller inlet at low speeds, as well as suppressing surge of the compression system at middle and high speeds. The flow recirculation process generated inside the casing treatment device removes the vortical structures at the impeller inlet tip section, therefore eliminating the regional rotating instability and stall. Besides, the casing treatment device turns the pressure rise characteristics of the impeller more negative by enhancing the impeller work input, but turns that of the vaned diffuser more positive by increasing the incidence at the vaned diffuser inlet. Combining both effects, mild surge and the deep surge are suppressed.

Flow instability evolution in high pressure ratio centrifugal compressor with vaned diffuser

Compressor stability is an essential issue for developing single-stage high pressure ratio centrifugal compressors. In this paper, a comprehensive stability map of a centrifugal compressor stage with the peak pressure ratio of 6.2 has been illustrated. Fast response pressure transducers are mounted inside the casing or the back-plate wall to measure internal transient behaviors. From Fourier analysis of pressure signals, various instabilities across time and length scales are identified. At low speeds, the impeller inlet rotating instability develops in scale and evolves into the inducer stall, which eventually induces the mild surge and deep surge of the compression system. At middle speeds, the compressor successively experiences stable state, mild surge, rotating instability, and deep surge. The mild surge region coincides with the dip region of the compressor S-shape pressure rise curve, and deep surge occurs when the compressor pressure rise raises to the second peak. At high speeds, mild surge and deep surge abruptly occur without preceding stall or rotating instability. To explain the complex surge behavior, the mechanical analogy between the compression system and the mass-spring-damper system is applied. Both mild surge and deep surge transients are found to belong to the dynamic instability of the mass-spring-damper system, and the exact form of the surge state will be determined by eigenvalues of the system governing equation.

In-pipe fibre optic pressure sensor array for hydraulic transient measurement with application to leak detection

Controlled hydraulic transient pressure waves can be used to actively assess the condition of pressurised pipelines (e.g. leak detection). One practical challenge to the application of the technique to buried water pipelines is in the pressure measurement – fast-response pressure transducers are costly and the deployment is limited to one sensor per access point (e.g. at air valves or fire hydrants). The measurement from a single sensor is always the superposition of the waves travelling upstream and downstream along a pipe. It is often difficult to determine the direction of the propagating waves, unless there are suitable locations nearby where additional measurements can be made. The research reported here develops a fibre optic pressure sensor array for in-pipe transient pressure measurement. The sensor array consists of five Fibre Bragg Gratings (FBGs), each formed into a custom designed pressure transducer. The cable containing the array has a diameter of approximately 4 mm, making it easy to insert into a pipeline through a single access point for distributed transient pressure measurement. The usefulness of the fibre optic pressure sensors for transient-based pipeline leak detection is validated through laboratory experiments. The measurement from each individual fibre optic sensor is generally consistent with conventional piezo-resistive pressure transducers over a wide frequency range. The sensor array configuration enables the determination of the directional information of waves, therefore enhancing the ability of pipeline defect detection and localisation.

High-frequency counter-flow plasma synthetic jet actuator and its application in suppression of supersonic flow separation

We come up with a control strategy for suppression of supersonic flow separation based on high-frequency Counter-flow Plasma Synthetic Jet Actuator (CPSJA). The main purpose of this investigation is to verify if its control authority can be enhanced by the jet/shock interaction. We use a blunt nose to generate a bow shock, a step on a flat plate to introduce a massive separation in a Mach 2 wind tunnel, and the CPSJA to generate Plasma Synthetic Jet (PSJ). In this study, pulsed capacitive discharge is provided for an array of CPSJAs, which makes the actuation (discharge) frequency f1 = 1 kHz, f2 = 2 kHz and f3 = 3 kHz. We use the high-speed schlieren imaging and fast response pressure transducers as well as a numerical simulation to investigate the quiescent PSJ properties, the interaction between the jet and bow shock, and its disturbance effect on the downstream separated region. The schlieren images show that PSJ is characterized by a succession of vortex rings; the jet strength weakens with the increase of frequency. A 4.5 mN jet thrust is found for all the frequencies. The simulation results show that jet/shock interaction produces vorticity in the vortex ring of the jet, enhancing turbulent mixing in PSJ so that a great deal of momentum is produced into the flow. We found the downstream flow is significantly disturbed by the enhanced actuation. Actuation with frequency of f2, f3 which is close to the natural frequency fn of the separation bubble suppresses the separation with the upstream laminar boundary layer being periodically attenuated, which has a better control effect than f1. The control effect is sensitive to the position where PSJ interacts with the shear layer, but the amount of energy deposited in one pulse is not crucial in a separation reduction in the experiment.

Stall Inception in a High-Speed Centrifugal Compressor During Speed Transients

The inception and evolution of rotating stall in a high-speed centrifugal compressor are characterized during speed transients. Experiments were performed in the single stage centrifugal compressor (SSCC) facility at Purdue University and include speed transients from subidle to full speed at different throttle settings while collecting transient performance data. Results show a substantial difference in the compressor transient performance for accelerations versus decelerations. This difference is associated with the heat transfer between the flow and the hardware. The heat transfer from the hardware to the flow during the decelerations locates the compressor operating condition closer to the surge line and results in a significant reduction in surge margin during decelerations. Additionally, data were acquired from fast-response pressure transducers along the impeller shroud, in the vaneless space, and along the diffuser passages. Two different patterns of flow instabilities, including mild surge and short-length-scale rotating stall, are observed during the decelerations. The instability starts with a small pressure perturbation at the impeller leading edge (LE) and quickly develops into a single-lobe rotating stall burst. The stall cell propagates in the direction opposite of impeller rotation at approximately one-third of the rotor speed. The rotating stall bursts are observed in both the impeller and diffuser, with the largest magnitudes near the diffuser throat. Furthermore, the flow instability develops into a continuous high frequency stall and remains in the fully developed stall condition.