High-frequency Pressure Measurements Research Articles

Plume and wall temperature impact on the subsonic aft-body flow of a generic space launcher geometry

Experimental and numerical simulation of launcher base flows are crucial for future launcher design. In experiments, exhaust plume simulation is often limited to cold or slightly heated gases. In numerical simulations, multi-species reactive flow is often neglected due to limited resources. The impact of these simplifications on the relevant flow features, compared to real flight scenarios, needs to be characterized in order to enhance the design process. Experimental and numerical investigations were carried out within the framework of the SFB/TRR 40 Collaborative Research Centre to study the impact of plume and wall temperature on the base flow of a generic small-scale launcher configuration. Wind tunnel tests were performed in the Hot Plume Testing Facility (HPTF) at DLR, Cologne, using subsonic ambient flow and pressurized air or hydrogen–oxygen combustion as exhaust gases. The tests were numerically rebuilt using the DLR TAU code employing a scale-resolved IDDES approach, including thermal coupling and detailed chemistry. The paper combines the experimental and numerical findings from the SFB/TRR 40 base flow studies and highlights the most prominent influences on the mean flow field, the pressure field, the dynamic wake flow motion, and the resulting aerodynamic forces on the nozzle. High-frequency pressure measurements, high-speed schlieren recordings, and time-resolved CFD results are evaluated using spectral and modal analysis of the one- and two-dimensional flow field data.

Flowfield and Noise Dynamics of Supersonic Rectangular Impinging Jets: Major versus Minor Axis Orientations

The current study explores the flowfield and noise characteristics of an ideally expanded supersonic (Mach 1.44) rectangular jet impinging on a flat surface. The existing literature is primarily concentrated on axisymmetric jets, known for their resonance dominance, pronounced unsteadiness, and acoustic signatures. In contrast, non-axisymmetric jets remain relatively less understood, particularly those impinging on a ground surface. By employing Schlieren imaging, high-frequency pressure measurements using high-bandwidth transducers, and particle image velocimetry (PIV), this research comprehensively examines the flow-acoustic phenomena. Schlieren imaging revealed distinct, coherent structures and strong acoustic waves, while pressure measurements at the impingement surface exhibited high-amplitude fluctuations, peaking at approximately 186 dB. Acoustic analysis identified multiple high-amplitude tones with unique directional characteristics, suggesting the potential for multiple simultaneous modes in rectangular jets. Furthermore, the PIV data elucidated differences in the jet shear layer and wall jet development attributed to the nozzle orientation. These findings contribute to a deeper understanding of non-axisymmetric jet behavior, offering insights relevant to fundamental flow physics and practical applications such as vertical takeoff and landing aircraft.

Unsteady Aerodynamics of the Retropropulsion Reentry Burn of Vertically Landing Launchers

During the vertical descent and landing of a launcher first stage with the aid of retropropulsion, commonly two main propulsive deceleration maneuvers are performed: the reentry burn in high altitudes at hypersonic to supersonic speeds and the landing burn shortly before touchdown at transonic to subsonic speeds. In the frame of the EU-funded H2020 project Retro Propulsion Assisted Landing Technologies (RETALT), the unsteady aerodynamics of those retropropulsion phases were studied. This paper presents results of experiments performed in the Hypersonic Wind Tunnel Cologne on the hypersonic reentry burn. The exhaust plume was simulated with pressurized air. Proper orthogonal decomposition was performed on high-speed schlieren videos, and spectral analyses of the time histories of the resulting modes were compared to the frequency content found in high-frequency pressure measurements. Dominant frequencies were found in the proper orthogonal decomposition modes for one and for three active engines. In the pressure measurements, dominant frequencies could only be observed for three active engines. The normalized pressure fluctuations are in the range of 0.002–0.012. Additionally, a good scaling of the pressures on the base area and in the wake of the configuration with the total pressure downstream of the bow shock could be confirmed, in the sense that the ratio of the local surface pressure to the total pressure downstream of the bow shock match for varying freestream Mach numbers.

Characterization and predictive modeling of a trajectory-oriented dual-mode scramjet combustor

Linear and nonlinear characteristics of a dual-mode scramjet combustor are investigated in ground-simulated acceleration and deceleration trajectory experiments. The experiments were conducted in the direct-connected transient flight trajectory simulator 1 at the Institute of Mechanics, Chinese Academy of Sciences. High-frequency pressure measurements, the schlieren, and CH* chemiluminescence high-speed imaging were applied for the diagnostics. Based on the quantitative analysis, the physical processes of acceleration and deceleration in general represent similar nonlinear characteristics. The linear characteristics are limited to the low-frequency oscillation period in the presence of physical governing mechanisms. A nonlinear predictive model of the dual-mode scramjet combustor based on historical measurements is proposed due to its generality for the acceleration and deceleration trajectory.

High temperature fuel impacts on combustion characteristics of a swirl-stabilized combustor

Future aircraft will require advanced thermal management systems to manage the increased heat loads resulting from the avionics, turbomachinery and other subsystems of higher performance, and high speed flight systems. Fuel is the primary coolant to alleviate these thermal management challenges. Heated fuel entering the combustor is likely to affect combustion performance due to dramatic changes in properties that impact fuel atomization, distribution, vaporization and fuel/air mixing. In this study, the combustion performance, emissions, and flame characteristics of a single-nozzle, swirl-stabilized combustor operating with high temperature fuels were investigated. Four fuels representing a wide range of jet fuel physical and chemical properties were studied at injection temperatures of 32, 121 and 260 °C. The test fuels included a conventional Jet A (A-2), a low cetane number alcohol-to-jet (ATJ) alternative fuel (C-1), n-dodecane (n-C12) and a high viscosity jet fuel blend (C-3A). Tests were conducted at a combustor pressure of 2 atm and inlet air temperature of 121 °C at global equivalence ratios (ϕ) of 0.10–0.24 for most fuel temperatures. Particulate matter (PM) emissions (concentrations and size distributions), and primary gaseous emissions were measured. High frequency pressure measurements and high-speed video were collected to assess fuel temperature impacts on combustion stability and flame characteristics, respectively. Results show that the heated fuel impacts on combustion were highly dependent on fuel type, temperature and ϕ. Qualitative results show that as fuel temperature increased, the flame in the primary zone appeared more compact, and less thermally radiant (lower soot). PM emissions were significantly impacted with reductions exceeding 95 % in particle number density for all fuels, and between 20 and 55 % reductions in mean particle diameters for aromatic-containing fuels. Gaseous emissions show up to ∼ 90 % reduction in unburned hydrocarbons (UHC) and up to 60 % reduction in carbon monoxide (CO), and corresponding moderate increases in NOx and CO2 with the heated fuel. The combustor acoustic response was slightly suppressed in frequencies of interest as fuel temperature increased. In general, the impacts of heated fuel on combustion were mostly beneficial and are primarily attributed to improved fuel atomization and reduced fuel evaporation timescales of the heated fuel. Specific impacts of fuel physical and chemical properties, and potential implications of heated fuel combustion are discussed.

Hysteresis and Bi-Stability in Transversely Excited Swirling Flows

Abstract We investigate the effect of transverse acoustic excitation on non-reacting swirling jets. The work is motivated by the azimuthal instabilities in annular gas turbine combustor which are one of the major challenges in aero-engines. We have designed and fabricated a multi-nozzle linear array combustor to simulate the flow conditions of an annular combustor. The nozzle features a dual co-rotating radial swirler configuration. Two compression drivers placed on either side of the combustion chamber are used to generate acoustic fields in the direction transverse to the flow. Simultaneous 2D PIV and high-frequency pressure measurements are conducted to measure the time-averaged velocity field and the chamber acoustics, respectively. It is observed that once the swirling jet is excited with a transverse acoustic forcing, it instantaneously transitions to a wall-jet state. In wall-jet state, the flow moves radially outwards and remains attached to the walls on either side of the nozzle, and is characterised by a recirculation zone with a strong negative axial velocity. In our experiments we demonstrate that transverse acoustic excitation can lead to a bistable state in swirling flows. We investigated the acoustic response to low-amplitude forcing on the combustion chamber by performing a forced acoustic response analysis using COMSOL. It is observed that acoustic forcing leads to a peak response of the radial and azimuthal velocities at the flare of the swirler, which could induce an axisymmetric deflection of the jet shear layer pushing it into a wall-jet state.

Detailed Measurement of Oxidizer-Rich Staged Combustion Injector Dynamics in Model Rocket Combustors

This paper presents experimental results from a model rocket combustor that uses a gas-centered swirl-coaxial injector element injecting warm oxygen and ambient temperature RP-2 fuel with 900 psia chamber pressure, representative of an oxidizer-rich staged combustion cycle. Results from the two injector geometries are analyzed and compared. One was a flat-faced injector with the element ending at a sudden expansion into the chamber. The other used a truncated cone at the injector exit: a design feature of specific interest that appears in drawings for the RD-170 main-chamber injector. To improve spatial resolution of high-frequency pressure measurements, the model injector was scaled up from RD-170 dimensions to facilitate measurements in the injector and head end of the combustor. Measured pressure spectra of the flat-face combustor showed a well-organized first axial mode and harmonics. The truncated-cone geometry showed a more diffuse spectra with higher-frequency modes, with a reduction in amplitude further upstream in the cone. Intermittency in measured dynamics is characterized via multiscale analysis. Adopting concepts from fractal theory, the stability margin of the injector configurations was quantified using the Hurst exponent of pressure oscillations. Comparatively, a larger range of fractal dimensions with smaller fluctuations in the initial reacting zone were observed for the flat-face injector.

Plume Stability During Direct Contact Condensation of Steam in a Crossflow of Subcooled Water, at High Mass Flux and With a Small Nozzle Diameter

Abstract The stability of a steam plume during direct-contact condensation into a crossflow of subcooled water is investigated for mass fluxes that are higher (>600 kg/m2s) and a nozzle diameter (2.4 mm) that is smaller than typically seen in the literature. The transition from a stable steam plume to an unstable plume associated with the formation and collapse of steam bubbles is characterized by high-speed imaging and high-frequency pressure measurements. Four regimes are observed: stable, condensation oscillation, transition, and unstable. A regime map and spectral signatures of the different flow regimes are provided. Results are compared with correlations from the literature, which are typically derived for lower mass fluxes, larger nozzles, and injection into stagnant pools of water.

Forced oscillation phenomenon of normal/oblique shock trains in a rectangular duct at Mach 2 and 3

The forced shock train oscillation is a common phenomenon in a practical working scramjet isolator. To evaluate the similarities and differences between the forced normal shock train oscillation and the forced oblique shock train oscillation, experiments were conducted in a rectangular duct at Mach 2 and 3. The normal/oblique shock trains were forced to oscillate for different back-pressure levels by changing the mechanical throttling ratio. High-frequency pressure measurements were mainly utilized for data acquisition and analysis. The frequency contents, the intermittency characteristics and the perturbation propagation of forced shock train oscillations were investigated in this work. Experimental results illustrate that the forced oscillation at Mach 3 is a composite pattern, which is dominated by the excited frequency and coupled with the inherent unsteadiness. This phenomenon is not obvious for the forced oscillation at Mach 2. For the forced oscillations at Mach 2 and 3, the maximum zero-crossing frequency is close to the downstream excitation frequency, and the whole intermittent region almost holds this value, which is very different from the self-excited oscillation pattern. By using the cross-spectral analysis, the frequency-dependent correlation was evaluated. It shows that the excitations and its harmonic contents can propagate upstream to affect the intermittent region and its downstream. Besides the excitation frequency and its harmonic contents, low-frequency contents (tens to more than a hundred Hz) can propagate nearby the separation region, and the propagation to downstream the separation bubble is very weak. Considering the time-lag between different transducers, the propagation speed of perturbations can be obtained. It implies that the propagation speed is independent of the magnitude of the back-pressure, but affected by the velocity of the incoming flow. Specifically, the higher the freestream Mach number is, the greater the propagation speed will be.

Singular spectrum analysis as a tool for early detection of centrifugal compressor flow instability

Centrifugal compressor machinery is subject to a potentially damaging phenomenon called surge at low mass flow rates. This effect may be preceded by a phenomena known as inlet recirculation — a flow reversal upstream of the impeller. A methodology to isolate inlet recirculation as a characteristic feature for monitoring of centrifugal compressor instability is presented in this study. The methodology is based on a nonparametric time series analysis technique called as singular spectrum analysis (SSA). SSA decomposes a signal into a number of Reconstructive Components (RCs), from which data trends and oscillatory components may be extracted. The frequency spectra of each RC and their relative contributions to the reconstruction of the original signal were examined and comparisons were made with spectral maps in existing literature. Individual and independent RCs were chosen to construct a compressor’s instability monitoring system. Additionally, the performance of SSA was determined by the Window length parameter. The effect of modification of this parameter was also studied, and the various viable choices of component for the basis of inlet recirculation diagnosis were considered. The methodology was implemented in pressure dynamical signals measured in an experimental centrifugal compressor rig. High frequency pressure measurements were taken at a number of flow conditions and locations within the compressor. The results demonstrated the potential of a methodology based on SSA to identify and extract oscillatory components with information about the local effect of inlet recirculation and eventually successfully monitor centrifugal compressor’s instability.

Experimental study on the forced oscillation of shock train in an isolator with background waves

The forced oscillation of shock train in an isolator with background waves have been investigated in a direct-connect wind tunnel. High-frequency pressure measurements were used to obtain the pressure histories in experiment and high-speed Schlieren technique was used to capture the shock train motion. Two types of oscillations during forced oscillation were studied and presented in detail. The large separation zone induced by leading shock is always located on one wall during type I oscillation, and switches between two walls during type II oscillation. Type II oscillation is a new type of oscillation that has not been observed in the forced oscillation occurring in an isolator with uniform incoming flow until now. During type II oscillation, the oscillation range of shock train is much larger than during type I oscillation, which makes type II oscillation have a greater impact on the performance and operating boundary of scramjet. The effect of backpressure oscillation amplitude and frequency on the two types of forced oscillation were studied. It was found that the backpressure oscillation amplitude and frequency have a significant effect on type I oscillation, but have little effect on type II oscillation. When the backpressure oscillation amplitude increases or the oscillation frequency decreases, the oscillation range of type I oscillation increases. If the backpressure oscillation amplitude is low or the oscillation frequency is high, type II oscillation does not exist. However, once type II oscillation occurs, its oscillation range does not change with the changes of backpressure oscillation amplitude and frequency. Finally, the different mechanisms of two oscillations were discussed. Type I oscillation is caused by backpressure perturbation and type II oscillation is caused by background waves.

Experimental study and analysis of shock train self-excited oscillation in an isolator with background waves

A study of shock train self-excited oscillation in an isolator with background waves was implemented through a wind tunnel experiment. Dynamic pressure data were captured by high-frequency pressure measurements and the flow field was recorded by the high-speed Schlieren technique. The shock train structure was mostly asymmetrical during self-excited oscillation, regardless of its oscillation mode. We found that the pressure discontinuity caused by background waves was responsible for the asymmetry. On the wall where the pressure at the leading edge of the shock train was lower, a large separation region formed and the shock train deflected toward to the other wall. The oscillation mode of the shock train was related to the change of wall pressure in the oscillation range of its leading edge. The oscillation range and oscillation intensity of the shock train leading edge were affected by the wall pressure gradient induced by background waves. When located in a negative pressure gradient region, the oscillation of the leading edge strengthened; when located in a positive pressure gradient region, the oscillation weakened. To find out the cause of self-excited oscillation, correlation and phase analyses were performed. The results indicated that the instability of the separation region induced by the leading shock was the source of perturbation that caused self-excited oscillation, regardless of the oscillation mode of the shock train.

Behavior and flow mechanism of shock train self-excited oscillation influenced by background waves

This study investigates shock train self-excited oscillation—influenced by background waves—occurring within an isolator in a direct-connect wind tunnel. The experimental data were obtained using high-speed Schlieren technique and high-frequency pressure measurements. Three oscillation modes of shock train self-excited oscillations were studied and have been presented in detail. These include Top-Large-Separation, Bottom-Large-Separation, and transition modes. The observed shock train with background waves conformed to one of the three modes, whereas the oscillation mode appeared randomly within a uniform incoming flow. To understand the differences of shock train unsteady behaviors in different modes, the distributions of intermittent region and zero-crossing frequency were compared. Furthermore, the mode and approximate location of the oscillations were judged by the distribution of wall pressure standard deviation. The influence of background waves on the shock train self-excited oscillation was then analyzed. For the shock train in a uniform incoming flow, the Strouhal number always lies in the range of 0.01–0.03, and the Strouhal number at both side walls is substantially equal. However, for the shock train influenced by background waves, the Strouhal number may fall outside the range of 0.01–0.03, and Strouhal number at the top and bottom walls is different. It was found that the wall pressure gradient caused by background waves affects the instability in the separation zone after the shock train leading edge, thereby influencing the unsteadiness of shock train self-excited oscillation.

Characterization of the Unsteady Aerodynamics of Optimized Turbine Blade Tips through Modal Decomposition Analysis

The present research focused on the analysis of the leakage flows developing from advanced blade tip geometries. The aerodynamic field of a contoured blade tip and of a high-performance rimmed blade were investigated against a baseline squealer rotor. Time-resolved numerical predictions were combined with high-frequency pressure measurements to characterize the tip leakage flow of each tip design. High spatial and temporal resolution measurements provided a detailed representation of the unsteady flow in the near-tip region and at the stage outlet. Numerical computations, based on the nonlinear harmonic method, were employed to assess the unsteady blade row interactions and identify the loss generation mechanisms depending on the tip design. The space- and time-resolved flow field was analysed by modal decomposition to identify the main periodicities of the near-tip and outlet flow and classify the most relevant sources of aerodynamic unsteadiness and entropy generation across the stage.

Two-dimensional unsteadiness map of oblique shock wave/boundary layer interaction with sidewalls

The low-frequency unsteadiness of oblique shock wave/boundary layer interactions (SBLIs) has been investigated using large-eddy simulation (LES) and high-frequency pressure measurements from experiments. Particular attention has been paid to off-centreline behaviour: the LES dataset was generated including sidewalls, and experimental pressure measurements were acquired across the entire span of the reflected shock foot. The datasets constitute the first maps of low-frequency unsteadiness in both streamwise and spanwise directions. The results reveal that significant low-frequency shock motion (with $St\approx 0.03$) occurs away from the centreline, along most of the central separation shock and in the corner regions. The most powerful low-frequency unsteadiness occurs off-centre, likely due to the separation shock being strengthened by shocks arising from the swept interactions on the sidewalls. Both simulation and experimental results exhibit asymmetry about the spanwise centre. In simulations, this may be attributed to a lack of statistical convergence; however, the fact that this is also seen in experiments is indicative that some SBLIs may exhibit some inherent asymmetry across the two spanwise halves of the separation bubble. There is also significant low-frequency power in the corner separations. The relation of the unsteadiness in the corner regions to that in the centre is investigated by means of two-point correlations: a key observation is that significant correlation does not extend across the attached flow channel between the central and corner separations.

A heterogeneous benchmark dataset for data analytics: Multiphase flow facility case study

Improvements in sensing, connectivity and computing technologies mean that industrial processes now generate data from a variety of disparate sources. Data may take a number of forms, from time-domain signals, sampled at various rates using a variety of sensors, to alarm and event logs. Novel techniques need to be developed to tackle the challenges of heterogeneous data. Testing such algorithms requires benchmark datasets that allow direct comparison of the performance of the methods. This work presents the PRONTO heterogeneous benchmark dataset. Experiments were conducted on a multiphase flow facility under various operational conditions with and without induced faults. Data were collected from heterogeneous sources, including process measurements, alarm records, high frequency ultrasonic flow and pressure measurements. The presented dataset is suitable for developing and validating algorithms for fault detection and diagnosis and data fusion concepts. Three algorithms are tested using the dataset, illustrating the applicability of the dataset.

Investigation of internal flow pattern of a multiphase axial pump

Due to recent advances in computational techniques, single phase performances of pumps can be easily estimated using CFD. However, when the working fluid is a mixture of liquid and gas (multi-phase flow) computations may struggle to capture the more complex physical phenomena that occurs, and the precision of the prediction worsen as the gas volume fraction at pump inlet is increased. To provide reliable benchmark for the improvement of simulations, in this study we experimentally investigated the internal flow pattern of a multiphase pump using high speed camera and high frequency pressure measurements, so as to obtain the data of pressure distribution inside of impeller, or to obtain blade loading, for several gas concentration conditions. Unsteady phenomena were identified at low flow rate conditions, and an explanation of their nature has been attempted.

Response of a Liquid Jet in a Multiple-Detonation Driven Crossflow

A bench-scale experiment was developed to investigate the response of a liquid injector site to the passage of multiple-detonation-type events with application to rotating detonation engines. The experimental apparatus consisted of two detonation channels providing sequential steep-fronted pressure waves perpendicular to the site of a plain orifice water injector. Detonation events were developed using a branched predetonator with ethylene and oxygen as propellants. The experiment was mounted within a pressure vessel for operation at both 1 and 10 bar chamber pressures. The effect of various injector geometries, chamber pressures, and detonation elevations from the injector were studied. Backlit high-speed videos provided insight into the transient response and breakup of the jet, and data were correlated based on jet-to-crossflow momentum ratios. High-frequency pressure measurements at the injector face provided the basis for a simple analysis for predicting recovery behavior using known parameters.

Analysis and modelling of unsteady shock train motions

The characteristics and mechanism for unsteady shock train motions were experimentally studied in a constant-area rectangular duct. High-speed Schlieren techniques and high-frequency pressure measurements were utilized in this research. The results show that the shock train undergoes periodical motions in response to downstream periodical excitations. The mechanism for unsteady shock train motions is that the shock train keeps changing its moving speed to change the relative Mach number ahead of shock train to match the varying back-pressure condition. It can be found that the unsteady shock train motion can be predicted well with a theoretical model, which is based on this mechanism. A correlation between the amplitude of shock train motions and some flow parameters was illustrated using an analytical equation, which was confirmed by the experimental results.

Experimental Study on Self-Excited and Forced Oscillations of an Oblique Shock Train

Two types of unsteady shock train motion, including self-excited oscillation and forced oscillation, were experimentally studied and compared in this work. The experiments were conducted in a rectangular isolator under Mach 3 incoming conditions. High-speed schlieren technology and high-frequency pressure measurements were adopted to do this research. The mean pressure distributions, the pressure standard deviation distributions, and the spatial frequency distributions were estimated for two types of unsteady shock train oscillations. It was found that the shock train may self-excite oscillations for Mach 3 incoming flow at relatively low frequencies (below 200 Hz) under steady backpressure conditions. With the downstream backpressure forcing, the shock train oscillated in a composite pattern, which was dominated by the imposed frequency content and coupled with the inherently self-excited frequency content. And, the self-excited oscillation was restrained, but was not eliminated, by the forced oscillation.