Unsteady Pressure Measurements Research Articles

Validation of a Methodology to Assess The Flutter Limit Cycle Oscillation Amplitude of Low-Pressure Turbine Bladed-Disks - Part I: Mach Number Effects

Abstract This paper presents a methodology to estimate the vibration amplitude of fluttering low-pressure turbine blades saturated due to friction effects. The study utilizes an analytical model that balances aerodynamic work and dry-friction work. The analytical predictions are compared against experimental results to validate the model. The first part of this paper focuses on the influence of the Mach number on the work balance between aerodynamic and mechanical components. It is observed that the vibration amplitude of low-pressure turbine rotor blades notably increases with higher Mach numbers. In addition, numerical simulations are employed to assess the influence of the Mach number on the critical damping ratio. The results demonstrate that an appropriate scaling of the critical damping ratio with the exit Mach number collapses all the damping versus inter-blade phase angle curves into a single curve. This finding validates the scaling of the aerodynamic damping for different pressure ratios. Unsteady pressure measurements were acquired, carefully post-processed to extract their flutter-induced peak components, and presented in a nodal diameter by nodal diameter basis. The post-processed data were then used to characterize the vibration amplitude observed in the experiments. The trends of the measured unsteady pressure on the casing of a rotating rig and the proposed model with the Mach number for different shaft speeds are in good agreement. The vibration amplitude and the mean unsteady pressure increase with the Mach number and exhibit a maximum with the shaft speed.

Mass and Heat Exchange in Rotating Compressor Cavities With Variable Cob Separation

Abstract Next generation aeroengines will operate at ever-increasing pressure ratios with smaller cores, where the control of blade-tip clearances across the flight cycle is an emerging design challenge. Such clearances are affected by the thermal expansion of the compressor disks that hold the blades, where acute thermal stresses govern operating life. The cavities formed by corotating disks feature a heated shroud at high radius and cooler cobs at low radius. A three-dimensional, unsteady and unstable flow structure is induced by destabilizing buoyancy forces. The radial distribution of disk temperature is driven by a conjugate heat transfer at Grashof numbers of order 1013. Such flows are further influenced by the heat and mass exchange with an axial throughflow of cooling air at low radius, where the interaction depends on the Rossby number and separation of the disk cobs. This paper is the first to study the effect of cob separation ratio on mass and heat exchange for compressor cavities. A model is developed to predict the cavity-throughflow interaction, and disk and fluid-core temperatures. The judicious use of a physics-based methodology provides reliable, reduced-order solutions to the complex conjugate problem, thereby making it appropriate for practical engine thermo-mechanical design. The model is validated by detailed experimental measurements using the Bath Compressor Cavity Rig, where variable disk cob spacings were investigated over a range of engine-representative conditions. The unsteady pressure measurements collected in the frame of reference of the rotating disks reveal new insight into the fundamentally aperiodic nature of the flow structure. This new understanding of heat transfer informs an expedient reduced-order model and enables more efficient design of future high pressure-ratio aeroengines.

Flow over two inline rough cylinders in the postcritical regime

This study investigates the flow behavior over roughened inline cylinders for postcritical flow, a parameter space with relatively little prior scrutiny. Two cylinders of the same relative surface roughness, ks/D=1.9×10−3, separated by a pitch (i.e., L, distance between the centers of two cylinders) between 1.175≤L/D≤10 are studied at Reynolds numbers from 3×105 to 6×105 using unsteady surface pressure measurements. As pitch ratio is increased from L/D=1.175, CD of the downstream cylinder increases sharply at (L/D)c=3.25. This critical pitch ratio (L/D)c is toward the lower end of the reported range for subcritical smooth cylinders. Asymmetric mean gap flow along with alternating reattachment is found for 1.5≤L/D<2.25 (i.e., two asymmetric modes in the gap, mode 1 and mode 2, that are the reflections of each other), and symmetric gap flow with a continuous reattachment is found for 2.25<L/D≤3. The gap flow is also symmetric for the closest pitch ratio tested of L/D=1.175. While the change in upstream cylinder drag coefficient with Reynolds number broadly follows that of an isolated cylinder, for the downstream cylinder, it is approximately independent. The critical separation is also insensitive to Reynolds number within 3×105≤Re≤6×105. Transitions between the reattachment and the co-shedding flow are predominantly continuous over the spanwise planes tested. On the other hand, alternating reattachment occurs in spanwise cells, where one sectional measurement exhibits the asymmetric mode 1 while a spanwise-adjacent section exhibits the asymmetric mode 2 or even symmetric flow. Previously reported maxima in the fluctuating lift and drag coefficients of the downstream cylinder at L/D≈2.4 at subcritical Reynolds numbers are absent in the current investigation.

Impact of Future Low-Emissions Combustor Technology on Acoustic Scaling Laws

A first-of-its-kind examination of broadband noise associated with a far-term advanced low-emission aerocombustor concept is presented. Because of design trends and expected cycle changes for future aircraft propulsion systems, noise generated by sources in the combustor are expected to become of increasing significance for airport-community noise. The paper assesses the impact on legacy semi-empirical noise-prediction methods from the expected radical departures from current combustor operating conditions and designs, such as fuel/air distribution and flame anchoring techniques. Such methods are essential in system-level noise assessments at the preliminary design stage for advanced air transports to assure that overall environmental goals are met. Detailed unsteady pressure measurements, obtained in a fundamental combustion noise experiment using a combustor rig at relevant pressures and temperatures are analyzed. In addition to an advanced far-term low-emissions concept, a reference configuration with the test section arranged to model a modern combustor sector was also studied. For the test rig in the current-generation configuration, the measured broadband acoustic data are reasonably well described by the acoustic-power scaling laws used in legacy semi-empirical noise-prediction methods. For the future-advanced configuration, the legacy scaling laws, with some notable exceptions, provide correct trends, but with much less accuracy.

Dynamics and spectral character of unsteady pressure field on afterbody of generic space launcher: Transonic flows

The unsteady pressure field over an axisymmetric backward-facing step was investigated experimentally at transonic freestream Mach numbers of 1.05, 1.2, and 1.4. The study was aimed at examining the influence of transonic freestream Mach numbers on the spatio-temporal character of the unsteady pressure field and on the dominant modes/mechanisms driving it. Surface flow visualization, schlieren, and unsteady pressure measurements were carried out as part of the experimental investigation. From oil flow visualization and schlieren, the reattachment region was identified, and consequently, the mean reattachment length was estimated. The mean reattachment length shows an increase with the increase in freestream Mach number. The coefficient of mean pressure along the rearbody imitates a classical backward-facing step flow profile and can be divided into three distinct regions. The peak values of the coefficient of mean pressure and the coefficient of root mean square of the fluctuation are seen to decrease with an increase in the freestream Mach number. Conventional spectral analysis reveals that as the freestream Mach number increases, the dominant peak in the spectra shifts to lower frequencies. From the spectra, three dominant fluid dynamic mechanisms depending on the freestream Mach number have been identified. Proper Orthogonal Decomposition (POD) analysis shows that 79–84 % of the total energy contribution comes from the first six modes. The temporal dynamics of the POD modes indicate three prominent mechanisms are responsible for the unsteady pressure field. Spectral analysis of POD modes indicates that the spectra are primarily driven by the first three POD modes for freestream Mach number of 1.05 and the first two modes for freestream Mach numbers of 1.2 and 1.4. Moreover, it reveals the presence of three dominant modes, and the freestream Mach number strongly dictates the dominant mode that is driving the pressure field.

Unsteady Surface Pressure Characterization on Launch Vehicle in Liftoff Configuration

Acquisition of unsteady pressure using remote scanners can provide practical advantages over making the same measurements directly using flush-mounted pressure transducers. However, pneumatic distortion of the signal is introduced by the tubing between the sensing module and surface port. This typically limits the usefulness of such measurements to averaged, steady data only, restricting the application of pressure scanners for unsteady pressure measurements in wind-tunnel and flight-testing applications. In this work, unsteady pressure on the surface of a 1.75%-scale model of the NASA Space Launch System Block 1B Cargo vehicle was remotely measured using an electronic scanning pressure module at the NASA Langley Research Center 14-by 22-Foot Subsonic Tunnel. Subsequent reconstruction using an empirically tuned, Wiener-filtered inverse transfer function produced excellent agreement compared to nearby surface-mounted transducers within a usable bandwidth of approximately 500 Hz. Interactions between the launch vehicle and the support tower produced strong oscillations that were absent in the vehicle-alone configuration, and it is likely that the wake interaction leads to a resonant flow response when the launch tower is immediately downstream of the vehicle.

Characterizing the Unsteady Flow Field in Low-Flow Turbine Operation

Abstract Current operational considerations require steam turbines to operate in a more flexible way, with more frequent and faster start-up and an increasing part-load operation. For very low mass flowrates, the interaction of highly separated flow with the high-speed rotor blades causes windage flow. This type of flow is characterized by increased temperature and highly unsteady flow, which forms vortex structures that rotate at a fraction of the rotor speed. If their magnitude is sufficiently high and the frequency is close to the blade eigenfrequency, non-synchronous vibration (NSV) can be induced. In this paper, low-flow turbine operation is investigated using a three-stage turbine rig that features an instrumentation concept focused on capturing aerodynamic and aeroelastic phenomena. Extensive steady probe, unsteady pressure, and tip-timing measurements are utilized. The experimental scope covers a wide range of operating points in terms of rotational speed and mass flowrates. Low-flow regimes are detected by a reversal in torque and increase in temperature. Unsteady measurements during transient operation identified large-scale vortical flow structures rotating along the circumference, so-called rotating instabilities (RIs). The onset, growth, and breakdown regimes of RI are characterized for different low-flow conditions. The quantitative characteristics of RI with regard to nodal diameter and rotational speed are derived by a cross-correlation of multiple unsteady sensors. The blade vibration measurements show a moderate structural response from unsteady aerodynamic excitation, indicating no significant NSV occurring in the present experimental setup. Later in the study, an acoustic excitation system has been applied to trigger a locked-in NSV without interrupting the coherent flow structures. From that, significant blade response has been observed, revealing a high degree of mistuning and damping of the rotor blading.

Aeroacoustic characteristics of a strut-braced high-lift device

The aerodynamic and aeroacoustic performance of a strut-based high-lift device were evaluated and demonstrated for six different strut models. The primary objective of the study was to investigate the impact of strut modifications on reducing noise levels. The aerodynamic characteristics are presented with the aid of surface pressure distribution on the airfoil that remained consistent across all the tested configurations. The aeroacoustic results are presented as the near-field surface pressure fluctuations and far-field noise measurements to attain a profound comprehension of the noise generation mechanism. Although the Albatros strut exhibited the greatest reduction in tonal noise, the directivity pattern and the overall sound pressure level of the radiated noise demonstrated that the medium height strut configuration can achieve noise reduction of up to 8 dB. The near-field unsteady surface pressure measurements are suggestive of harmonic oscillations. The coherence studies carried out have shown a decrease in the tonal coherence for the small height strut configuration while the velocity field measurements performed in the wake of the high-lift device show no significant variation in flow patterns between different strut configurations.

Indexes for evaluation of dynamic characteristics of pressure-sensitive paint based on pressure sensitivity and frequency response

Two evaluation indexes were proposed to select the optimal PSP for unsteady pressure measurement from various PSPs with different pressure sensitivity and frequency response. An effective sensitivity coefficient calculated by pressure sensitivity and gain attenuation due to the response delay was proposed. Furthermore, an effective amount of intensity change was recommended, which takes into account the emission intensity and the effective sensitivity coefficient, because the magnitude of the intensity of a PSP is also important for unsteady-pressure measurement with high-speed sampling. A total of five types of PC-PSPs developed in previous studies were compared: two types of fast-response PC-PSPs using Pt(II) meso-tetra (pentafluorophenyl) porphine (PtTFPP) with poly(isobutyl methacrylate) (poly(IBM)) and ruthenium complex with RTV silicone, respectively, and three types of PC-PSPs using PtTFPP with poly[1-trimethylsilyl)-1-propyne] (poly(TMSP)). A comparison was made using the proposed evaluation indexes under various pressure ranges. The results shows that poly(TMSP)-based PC-PSP has a high effective sensitivity coefficient at pressures less than 20 kPa. On the other hand, poly(IBM)-based PC-PSP has the highest effective sensitivity coefficient at a pressure of 100 kPa. The effective amount of intensity change of poly(TMSP)-based PC-PSP is the highest at 2 kPa, but that of poly(IBM)-based PC-PSP is the highest at a pressure higher than 5 kPa among the evaluated PC-PSPs due to its high luminescence intensity. A PSP with high emission intensity will provide high performance in terms of fluctuation of emission intensity detected by the photodetector when the excitation intensity and the exposure time are limited due to limitations of optical equipment or high-speed sampling.

Unsteady loads on circular cylinder in cross-flow: An observation in natural wind and comparison with wind tunnel experiments

We present the synchronized unsteady wall pressure measurements performed at the circumference of a full scale 35 m circular chimney submitted to natural wind. The atmospheric turbulent boundary layer generates a turbulence intensity of 15 % at the altitude of the measurements and the corresponding Reynolds number is 1.13 106. For the first time, the steady and unsteady pressure distributions around the circumference are presented, despite large difficulties in the measuring process. The analysis of the pressure distribution using the bi-orthogonal decomposition (BOD) is a key factor in the success of the processing. By comparison with wind tunnel similar measurements, with 2D cylinders, the time-averaged pressure distribution is found between the one of a subcritical regime and a supercritical regime, slightly closer to the subcritical regime notably because the base pressure coefficients are similar. The atmospheric turbulence is an essential ingredient for the unsteady pressure response on the chimney, which generates high levels of RMS pressure, much higher than all those measured in wind tunnel, even under the turbulent flow generated by a grid. The analysis by BOD allowed the extraction of a term which is mainly generated by alternate vortex shedding. Despite a very noisy signal, the Strouhal number is determined to be 0.20–0.24. The corresponding shape of the pressure distribution is compared with those found in wind tunnel: it is shown that the full scale field measurements are close, but not identical, to the small scale measurements in wind tunnel on a rough cylinder submitted to a grid generated turbulent flow.

Nonembedded measurement method based on amplitude correction for unsteady surface pressure estimation in a high-subsonic compressor cascade

Unsteady surface pressure (USP) is a crucial physical quantity for studying blade loads and boundary layer flow. However, the unique characteristics of high-load compressor cascades, including thin walls, narrow passages, and significant geometric deflection angles, pose considerable challenges for USP measurements. The primary innovation of this study lies in proposing a nonembedded measurement method based on amplitude correction for USP measurement, which can provide frequency and amplitude information of dynamic pressure on measured surface. Unlike existing methods that rely on theoretical models for amplitude correction, the proposed method establishes amplitude correction transfer models using Gaussian process regression (GPR) based on pre-calibration data. This innovation overcomes the constraints faced by theoretical models when dealing with high-speed complex geometries, such as those found in compressor blades, which are limited by geometric and Reynolds number considerations. The principle of this method is to perform indirect measurement by leading the airflow out of the measured surface through an air tube and correcting the amplitudes through transfer models. An acoustic calibration system was developed for amplitude calibration and measurement verification. The results demonstrate that the developed method accurately provides the dynamic pressure frequencies in the range of 1–10 kHz with a tube length of ≤ 55 mm and from 1 to 7 kHz with the tube length of 80 mm. Verification results proof that the GPR statistical model can reasonably estimate amplitude attenuation. Ultimately, the developed method was successfully applied to unsteady pressure measurement in a high-subsonic compressor cascade and showed highly encouraging results.

Simultaneous measurement of unsteady surface pressure and leading edge noise of a NACA0012 airfoil

Airfoil-turbulence interaction noise, which is created whenever turbulent flow encounters an airfoil, is a major source of unwanted noise emitted byaircraft and turbomachinery. Despite being a major research focus overthepast decades, the accurate prediction and understanding of airfoil–turbulence interaction noise remain an open question. This experimental investigation involved simultaneous measurements of unsteady pressure fluctuations at 61 locations across the surface of a NACA0012 airfoil and far field sound pressure at a phased 64-microphone array. Various square-mesh turbulence grids were placed upstream of the airfoil to vary the isotropic free stream turbulence. Additionally, the influence of the airfoil’s angle of attack as well as different mean flow speeds on the unsteady surface pressure and radiated far field noise was investigated. Ultimately, the links between the simultaneously acquired unsteady surface pressure on the airfoil and the microphone array sound pressure were studied. This provided insight into the mechanisms of leading edge noise source generation and their influence on the radiated far field noise.

Tandem cylinder aeroacoustic sources

Tandem cylinders in turbulent flow are found in aircraft landing gear, chimney stacks, power lines, and bridge piers. The aim of this paper is to present detailed unsteady surface pressure measurements for tandem cylinders. These unsteady surface pressures are the major sources of noise in this important fundamental test case. A series of experiments were conducted in the UNSW anechoic wind tunnel to investigate the sound generation and flow distortion from a downstream cylinder interacting with wake from an upstream cylinder. This wake is highly anisotropic, containing large vortex shedding scales as well as energetic broadband components. The experiments were conducted using PIV (Particle Image Velocimetry) to obtain instantaneous flow fields and remote microphones to capture unsteady surface pressure measurements. Microphones were positioned outside the flow to record far-field noise. The cylinder was rotated to obtain surface pressure around the circumference. The acoustic results for the tandem cylinder case were compared with noise generation from a single cylinder in uniform flow. An increase in both surface pressure and far-field noise was observed in the tonal and broadband components. It was concluded that interaction with the wake of the upstream cylinder significantly increased the unsteady surface pressure of the downstream cylinder.

On the flow unsteadiness and operational characteristics of a novel supersonic fluidic oscillator

A novel supersonic jet oscillating method is investigated both experimentally and numerically. A rectangular primary supersonic jet is issued into a confined chamber with sudden enlargement. Secondary control jets are issued from the top and bottom backward-facing step regions formed due to sudden enlargement. The primary jet is oscillated in the transverse direction by blowing the secondary jets in the streamwise direction in a pulsating manner with a phase shift. The out-of-phase secondary jet blowing causes the primary jet to periodically adhere to the upper and lower part of the confined chamber, causing flapping of the primary jet and acting as a supersonic fluidic oscillator. The supersonic jet oscillation characteristics are experimentally investigated using shadowgraph type flow visualization technique and steady and unsteady pressure measurements. Quantitative analysis of the shadowgraph images using the construction of y – t and y – f plots reveals the presence of periodic jet oscillation with a discrete dominant frequency similar to the secondary jet excitation frequency. The existence of linearity between the excitation frequency and the flapping jet frequency on the low-frequency (0.66–6.6 Hz) side is first proven experimentally. Later, the high-frequency (16.67–5000 Hz) operation extent of the supersonic fluidic oscillator is further demonstrated using unsteady computational studies owing to the existing experimental facility's limitations. A reduced-order analytical framework has also been proposed to investigate the limiting oscillation frequency. It is found that the limiting frequency predicted from the proposed analytical model shows fairly good agreement with the computationally predicted results (5 kHz).

Flow Field Over Boat-Tail Region of Heat Shield of a Typical Launch Vehicle at Mach 2

Boat-tailed heat shields offer a higher payload volume but lead to flow separation and associated effects. In the present paper an attempt is made to find out the effect of boat-tail angle on the over all flow field with special emphasis on acoustic load near the boat-tail region. Experiments and CFD routes are used to investigate the flow field. Experiments consist of steady and unsteady pressure measurements and flow visualisation through schlieren and oil flow. CFD simulations are also carried out for the wind tunnel conditions. Results indicate that fluctuating pressure co-efficient increases with flow separation, peaks at reattachment point and reduces further downstream. An Increase in boat-tail angle from 20" to 30" leads to about I2Vo increase in acoustic load-s. All these studies were carried out at a Mach number of 2 and a Reynolds number of 4 x t}o based on model length.

Pressure Fluctuations in the Base of a Typical Re-Entry Module

Unsteady pressure measurements were carried out on the flow side and around the base of a blunt-nosed bluff body representing a spacecraft re-entry module in the Mach number range of 0.8 to 4.0. The results show that the pressure fluctuations in the base region are much higher than in the attached flow region of the body. Most of the energy in the pressure fluctuations in the base region is in the low-frequency band, while fluctuations in the attached flow region occur over a wider bandwidth. In the separated flow region at the base, highest pressure fluctuations were observed around the rear stagnation point. The frequency spectra do not show any dominant tone attributable to the configuration at any measurement location. The observations suggest that vortex shedding behind such 3-d bluff bodies could be relatively weak and that most of the fluctuating energy in the separated flow region could be due to random oscillations of the rear stagnation point.

Supersonic Flow Separation in Front of a Forward-Facing Step

A detailed experimental study has been done to understand the Shock Wave-Boundary Layer Interaction (SWBLI) over a Forward-Facing Step (FFS) of height (h) in a Mach 2.5 flow using Particle Image Velocimetry (PIV). In addition to PIV, unsteady wall pressure measurements, high-speed schlieren and surface oil flow visualization have been studied. Flow separation using surface oil flow visualization was observed around 4.1 step heights upstream of the step face. Shock motions extracted from the schlieren image sequence shows a peak frequency of approximately 1000 Hz, which corresponds to the dominant (low) frequency at which the shock oscillates; this being two orders in magnitude smaller than the characteristics frequency of the incoming boundary layer. Also, spectra of the unsteady wall pressure data shows that the shock oscillates with a peak frequency of approximately 1000 Hz. The effect of shock location (xs) from instantaneous velocity PIV fields in the cross-stream plane due to an upstream influence parameter (line-averaged boundary layer streamwise velocity, ul) and a downstream parameter (separation bubble or recirculation region area, A) is studied. Instantaneous cross-stream PIV measurements show significant variations of the separation shock location, incoming boundary layer velocity fluctuations and recirculation region. Besides, conditionally averaged velocity field based on incoming boundary layer streamwise velocity fluctuations and separation bubble area in the cross-stream plane shows that the shock location (xs) is correlated to the size of the separation bubble and not to boundary layer velocity fluctuations.

Effect of Protrusions on the Surface Flow and Pressure Fluctuations in the Inter-stage Region of A Launch Vehicle Model

Experiments were carried out to assess the effects of geometrically scaled and viscous scaled protrusions on the surface flow and unsteady pressures in the inter-stage region of a typical launch vehicle model in the Mach number range 0.8 to 1.6. Formation of bow-waves due to the protrusions, and three-dimensional boundary layer separations and reattachments of the shear layer were seen from flow visualizations. Measurements of unsteady pressures show that protrusions cause substantial increases in fluctuation levels on the core vehicle. The highest levels on fluctuations were found to be at a location behind the viscous scaled protrusion, close to the shear layer reattachment. The spectra of pressure signals suggest that the reattachment may be periodic in the range 0.90 £ M £ 1.60; the associated frequency was found to vary in the range 1 kHz to 3.2 kHz (corresponding to Strouhal number in the range 0.034 to 0.051) depending on M. Streamlining the viscous-scaled protrusion was found to replace bow-waves with multiple oblique shocks, resulting in reduced levels of pressure fluctuations, comparable to the basic levels.

Control of Back-pressure During Wind Tunnel Testing of A Two Dimensional Ramjet Air-intake Model

For optimum utilization of available run-time in blowdown wind tunnels, the controls, instrumentation, data acquisition and various other related systems associated with measurements must be efficiently synchronized. Testing of the air-intake model is a typical example where large number of steady and unsteady pressure measurements need to be combined with control of back pressure at the exit plane of the intake. This paper presents the development of a PC-based control system for testing a ramjet air-intake model at supersonic Mach numbers in the 0.3m trisonic wind tunnel at CSIR-National Aerospace Laboratories, Bangalore. The back pressure of the two-dimensional ramjet air intake model was controlled using a rectangular butterfly valve driven by a stepper motor external to the test section. Commands to vary the valve position were programmed in a PC and the valve position acquired from a potentiometer. The motor control and pressure data acquisition were operated in a hand-shake mode to enable hands-off execution of tests and maximize the data throughput from each blowdown.

Experimental investigation of flow unsteadiness over four types of spiked-blunt body at Mach 2.2

Flow unsteadiness over four types of spiked-blunt body are investigated through wind tunnel experiments at a freestream Mach number of 2.2 at 0° and 4° angle of attack. The flow pulsation is visualized with high spatiotemporal resolution schlieren and then the shock-related unsteadiness is investigated through proper orthogonal decomposition (POD) method. Firstly, two special states in the suppressed pulsation mode are observed, namely the phase of withhold and the phase of smash. Flow over the standard aerodome spike will continuously transform between the two states, which can be distinguished from the distinct flow structures. Spectral analysis of unsteady pressure measurement detected the dominant frequencies of 2.2 kHz and 3.2 kHz for pointed spike and standard aerodome spike, respectively, and POD analysis further finds that the signal originates from the periodic motion of the aftershock. The large-amplitude flow pulsation (amplitude 0.23D) in typical pulsation mode is transformed into the high-frequency small-amplitude pulsation (amplitude 0.09D) in the suppressed mode. In addition, it is found that the variation of angle of attack makes little influence on the dominant frequency. And it is hard to find the main frequency of double/inverse aerodome spike, which indicates that the flow unsteadiness of different aerodome cases is not the same.