Indraft Wind Tunnel Research Articles

Influence of combinational distortion generator geometrical parameters on the steady and dynamic components of the total pressure distortion

To construct a method for designing a total pressure distortion generator with adjustable, steady, and dynamic components, we experimentally investigated the influences of combinational distortion generator geometrical parameters on the steady and dynamic components of the total pressure distortion based on an indraft wind tunnel. Different types of annular plates and cylindrical rods were installed upstream of a traditional baffle total pressure distortion generator; the total pressure distortion was then measured using rotatable total pressure rakes. The installation of cylindrical rods upstream of the baffle plate could maximally reduce the ratio of steady and dynamic components of total pressure distortion by 30%. Increasing the number and height of the rods, as well as the axial distance between the rod and baffle plate, were effective methods of reducing the steady components of total pressure distortion. Similarly, changing the geometrical parameters of the annular plate was also an effective method for adjusting the ratio of steady and dynamic components of total pressure distortion. The ratio of the steady and dynamic components of total pressure distortion first experienced a decrease and then increased variance law with the increase in the height and circumferential scale of the annular plate. During this procedure, the steady and dynamic components ratio of total pressure distortion could be maximally reduced by 24% and maximally increased by 20%.

Pressure Losses Downstream of a Compact Valve in the Inlet Chamber of an Intermediate-Pressure Steam Turbine

Deep knowledge about pressure and energy losses in each part of a steam turbine is crucial for assuring the required efficiency and operational reliability. This paper presents the experimental as well as the numerical study of pressure losses in the inlet chamber of an intermediate-pressure steam turbine. Measurements were carried out on a complex model, where not only was there an inlet turbine chamber, but also a compact valve assembly situated upstream and nozzles situated downstream. The compact valve as well as the turbine inlet chamber were relatively small. Therefore, greater pressure losses were expected. The aerodynamic laboratory of the Institute of Thermomechanics of the Czech Academy of Sciences was responsible for acquiring the measurements, which were carried out in a modular in-draft wind tunnel. In order to learn further details, numerical simulations were carried out. Doosan Skoda Power was responsible for this. A package of ANSYS software tools was used. Measured data were described and compared with numerical ones. Pressure losses were generalized in the form of the total pressure loss coefficient. As a result, pressure losses in similar turbine compact inlet chambers can be predicted with the required accuracy.

Study on Losses in the Intermediate-Pressure Steam Turbine Inlet Chamber

Know-how about energy and pressure losses in all steam turbine parts is crucial to guarantee enhanced operational reliability and efficiency. This paper focuses on studying pressure losses in the intermediate-pressure turbine inlet chamber. Measurements were performed on a complex model consisting of the turbine inlet chamber, a compact valve assembly situated upstream, and nozzles located downstream. These intermediate-pressure turbine parts are larger due to the greater volumetric mass flow than the high-pressure turbine parts. However, their inner parts are proportionately smaller, which causes greater pressure losses. Measurements were carried out in the Aerodynamic laboratory of the Institute of Thermomechanics of the Czech Academy of Sciences, where the model was installed in an in-draft wind tunnel. The results were complemented by numerical simulations performed in the Doosan Skoda Power company using ANSYS software tools. Pressure losses were evaluated using the total pressure loss coefficient and, as a result, can be predicted in similar turbine inlet chambers with the required accuracy.

Application of laser energy deposition to improve performance for high speed intakes

Research interest has been growing in recent years in supersonic transport, particularly supersonic propulsion systems. A key component of a commonly studied propulsion system, ramjets, is the air intake. For supersonic propulsion systems a major factor in the overall efficiency is the intake pressure recovery. This refers to the ratio of the average total pressure after the intake to that of the freestream. One phenomenon that can have a large effect on this performance index is flow separation at the inlet. The aim of this work is to examine how pulsed laser energy deposition can be used to improve pressure recovery performance by reducing flow separation at the inlet. This research examines the effects of pulsed laser energy deposition upstream of an intake with an axisymmetric centrebody in a Mach 1.92 indraft wind tunnel. Laser frequency was varied between 1 and 60 kHz with an energy per pulse of 5.6 mJ. Schlieren photography was used to examine the fundamental fluid dynamics while total and static pressure downstream of the intake diffuser were measured to examine the resulting effect on the performance. Schlieren imaging shows that the interaction between the laser generated thermal bubble and the leading edge shock produced by the centrebody results in a significant reduction in separation along the intake cone. Analysis of the schlieren results and the pressure results in tandem illustrate that the average separation location along the length of the centrebody directly correlates to the pressure recovery observed in the intake. At the optimal laser frequency, found for this Mach number to be 10 kHz, the pressure recovery is found to increase by up to 4.7%. When the laser power added to the system is considered, this results in an overall increase in propulsive power of 2.47%.

Repetitive Energy Deposition at a Supersonic Intake in Subcritical and Buzz Modes

This study experimentally investigated the effects of repetitive laser energy deposition using a supersonic intake model with a central conical compression surface and a cross-sectional duct in a Mach 1.92 indraft wind tunnel: especially in the subcritical and buzz modes. A single-pulse energy deposition was observed to suppress the flow separation at the compression surface by “sweeping” a shock wave system with the thermal bubble generated by the energy deposition. The duration of the sweeping effect was approximately in the subcritical mode and in the buzz mode. Furthermore, repetitive deposition of laser pulse energies was observed to moderate instabilities in both modes, and it increased the pressure recovery by as much as 8%; also, the occurrence of buzz was delayed, thereby widening the stable, subcritical regime. In both modes, there was a threshold value for the laser pulse repetition frequency , which corresponded to the duration of the sweeping effect (e.g., in subcritical mode). Below this frequency, the increase in the pressure was proportional to the repetition frequency; whereas above , the effect per single pulse was reduced.

Experimental Investigation of the Boundary Layer Transition on a Laminar Airfoil Using Infrared Thermography

The transition in the boundary layer is investigated using infrared thermography (IRT). The study is carried out on a laminar airfoil in the transonic intermittent in-draft wind tunnel. The transition in the boundary layer is evocated using transition-generator strips of different thicknesses at two Mach numbers: 0.4 and 0.8. The tested transition-generators thickness to boundary layer displacement thickness ratio was from 0.42 to 1.25. The Reynolds number respect to the airfoil chord is: Re = 1.0 – 1.4·106. The six cases for different transition-generators thickness ratios were compared. The behaviours of laminar and turbulent boundary layers are discussed. The use of IRT has been proven to be an appropriate tool for detecting the transition of the boundary layer in high-speed wind tunnel testing.

A detailed statistical study of unsteady wake dynamics from automotive bluff bodies

There is a growing emphasis from automotive manufacturers on efficient performance, with a greater need to leverage unsteady aerodynamics data in the design process. Simplified bluff bodies are regularly used to generate flow features which replicate specific behaviour of flow around real vehicles. The unsteady aerodynamics of a squareback and a 40∘ slant back Windsor body was investigated here using Computational Fluid Dynamics (CFD) at Re = 2.56 × 106. The results were generated using a Lattice-Boltzmann-based Very Large-Eddy Simulation (VLES) solver and experimental measurements were taken in a static-ground, in-draft wind-tunnel. This paper compares unsteady force coefficients, base surface pressures and vortex interaction in the wake. Statistical methods were defined and applied to identify the correlation present in the sampled signals, propose improvements to the calculation of confidence intervals and understand interaction between different coherent flow features. The slant modification was found to significantly shorten the integral time-scales in the near-wake. Spectral analysis of point probes revealed that the 40∘ slant modification favoured higher frequency content. Independent, modified confidence intervals showed a log-linear variation with sample sizes.

Spatial density fluctuation of supersonic flow over a backward-facing step measured by nano-tracer planar laser scattering

Spatial density fluctuation of supersonic flow over a backward-facing step is studied by nano-tracer planar laser scattering, aero-optic method and proper orthogonal decomposition analysis. Measurements were carried out in a Ma = 3.0 low-noise indraft wind tunnel. By varying the superficial roughness of the wall upstream from the step, supersonic laminar flow and supersonic turbulent flow were generated over a backward-facing step. The reattachment regions of these two flows were focused on. Flow structures of supersonic boundary layer, free shear layer, $$Kelvin - Helmholtz$$ vortices and reattachment shock were revealed. Based on the density-grayscale calibration of the flow field, the density distribution and optical path difference were calculated. Proper orthogonal decomposition was applied to 400 samples of optical path difference distributions which were generated by performing ray tracing method. According to the principle of proper orthogonal decomposition and the aero-optic basic equations, peaks in the first mode of proper orthogonal decomposition could reveal intensive spatial density fluctuations. The results showed that around the location of the reattachment point, density changed dramatically in both of two flows spatially. The location of reattachment point judged by the intensive density variation basically agreed with that judged by velocity field or flow structures. Thus, it was found that reattachment of supersonic backward-facing step occurred with intensive spatial density fluctuation which caused complex aero-optic aberration.

円錐スパイクとエネルギー付加の組み合わせによる鈍頭物体の 超音速抗力低減

Drag reduction performance over a blunt-body by a combination of a conical spike and repetitive energy depositions is investigated. Experiments are conducted in a Mach 2 supersonic flow in an in-draft wind tunnel, varying the length and the apex angle of the spike on a flat head body and the repetitive frequency of laser pulse energy depositions from 0 to 60kHz. Computational fluid dynamics calculations are done to diagnose the flow fields. Comparison is made between the experiment and calculation with respect to visualized flow fields and to the drag. When the length of the spike is shorter than the shock stand-off distance over the body without the spike, the effect of the spike is significant in low repetitive frequencies of up to 40kHz. In even higher repetitive frequencies, the impact of the spike becomes weak; the drag reduction is primarily attributed to energy deposition.

Effectiveness of jet location on mixing characteristics inside a cavity in supersonic flow

The incorporation of cavities within supersonic combustion chambers is an effective means of slowing down the flow for fuel injection and consequent stable combustion. Understanding the flow physics associated with such flows, especially with the injection of a gas to replicate fuel injection, are essential for the optimum design of supersonic propulsion mechanisms. An experimental investigation was performed on a rectangular open cavity with upstream injection model in a Mach number of 1.9 using a trisonic indraft wind tunnel. A rectangular open cavity of dimensions L/D=5, 100mm in length (L) and 20mm deep (D), was adopted, and it was embedded into the lower wall of the test section. An air jet with a jet-to-freestream momentum flux ratio of J=1.2, 2.7 and 5.3 was injected upstream of the cavity. To evaluate the effect on mixing and flow stability the jet position, measured from the front edge of the cavity, was varied between 0.1L and 1L. The flow field was visualized using schlieren photography, particle image velocimetry, and oil flow measurements. It is found that the mixing characteristic within the cavity when the jet is positioned 0.1L is enhanced independent on the J value because the turbulence intensity of the flow velocity within the cavity is strongly influenced by the jet interaction which lifted the flow from the floor of the cavity compared to the other jet positions. However, the flow over the cavity is unstable at all jet positions. The separation shock formed at the front edge of the cavity oscillates significantly for the case where the jet is located at 0.1L because the separation shock location coincides with the compression shock behind the jet.

Flow visualization of supersonic laminar flow over a backward-facing step via NPLS

An experimental study on a supersonic laminar flow over a backward-facing step of 5 mm height was undertaken in a low-noise indraft wind tunnel. To investigate the fine structures of Ma = 3.0 and 3.8 laminar flow over a backward-facing step, nanotracer planar laser scattering was adopted for flow visualization. Flow structures, including supersonic laminar boundary layer, separation, reattachment, redeveloping turbulent boundary layer, expansion wave fan and reattachment shock, were revealed in the transient flow fields. In the Ma = 3.0 BFS (backward-facing step) flow, by measuring four typical regions, it could be found that the emergence of weak shock waves was related to the K–H (Kelvin–Helmholtz) vortex which appeared in the free shear layer and that the convergence of these waves into a reattachment shock was distinct. Based on large numbers of measurements, the structure of time-averaging flow field could be gained. Reattachment occurred at the location downstream from the step, about 7–7.5 h distance. After reattachment, the recovery boundary layer developed into turbulence quickly and its thickness increased at an angle of 4.6°. At the location of X = 14h, the redeveloping boundary layer was about ten times thicker than its original thickness, but it still had not changed into fully developed turbulence. However, in the Ma = 3.8 flow, the emergence of weak shock waves could be seen seldom, due to the decrease of expansion. The reattachment point was thought to be near X = 15h according to the averaging result. The reattachment shock was not legible, which meant the expansion and compression effects were not intensive.

An experimental study on fine structures of supersonic laminar/turbulent flow over a backward-facing step based on NPLS

Fine structures of supersonic flow over a 5 mm high backward facing step (BFS), including expansion wave fan, reattachment shock, supersonic boundary layer were measured in a Ma=3.0 low-noise indraft wind tunnel. By varying the superficial roughness of the wall upstream from the step, supersonic laminar flow and supersonic turbulent flow could be formed over a BFS. Measurements on the spatiotemporal features of the holistic flow field and the fine structures in four typical regions were carried out using NPLS (nano-based planar laser scattering). Flow structures, including expansion wave fan, reattachment shock, supersonic boundary layer and its separation, reattachment and redevelopment are revealed by measuring the holistic structure of the transient flow field. Comparing the two time-averaged flow fields with each other, it is apparent that supersonic turbulent flow over a BFS (STF-BFS) has a larger expansion angle and a shorter recirculation region, and its redeveloped boundary layer increases at a smaller obliquity while the angle of reattachment shock is the same for the supersonic laminar flow over a BFS (SLF-BFS). With regard to time-evolution features, the K-H vortices in the SLF-BFS suffers from shearing, expansion, reattachment and three-dimensional effects while in the STF-BFS large-scale structures are affected by the incline and distortion at the reattachment point due to expansion, viscosity and reverse-pressure. Studies on local regions indicate that in the SLF-BFS, the emergence of compression waves which distinctly converge into a reattachment shock is due to the local convective Mach number and the inducement of K-H vortices in the free shear layer. Nevertheless, in the STF-BFS, compression waves and K-H vortices are barely evident, and the formation of a reattachment shock is related to the wall compressive effect.

Supersonic Drag Reduction with Repetitive Laser Pulses Through a Blunt Body

A drag over a flat-nosed cylinder with repetitive laser pulse irradiations ahead of it were experimentally measured in a Mach-1.92, in-draft wind tunnel. Laser pulses were focused using a plane-convex lens fabricated on the nose of the cylinder at a repetition frequency of up to 10 kHz and power of 70 W at a maximum. The drag was measured using a low-friction piston which was backed by a load cell in a cavity at a controlled pressure. Under the experimentally available operation conditions, an up-to-3 % drag reduction and an efficiency of energy deposition of about 10 were obtained. Mechanisms of the drag reduction are analyzed based on experimental flow visualization and numerical simulation.

Experimental Study on Fast Ionization Wave in Supersonic Flow

Results of a supersonic flow ionization experiment are described in this paper. In the experiment, the Mach 3 in-draft supersonic wind tunnel and the air were used. As an ionization method, the first ionization wave (FIW) was used because FIW is expected to produce a large number of electrons and as a result, to enhance an electrical conductivity of the boundary layer on a wall of the wind tunnel. The main objective of the experiment was to investigate discharge formation inside the boundary layer by FIW. A repetitive power supply was designed, built-up, and tested. The power supply has shown to provide a peak voltage of ±32kV, a pulse width of 100ns, a rapid voltage rise time of 0.6kV/ns and a repetition rate of about 1kHz. When the high voltage pulse was applied, a bright light emission from the air was observed near the wall, and it extended downstream over a long distance. The propagation of FIW along the flow was examined using two capacitive voltage probes. As a result, the propagation of FIW was confirmed, and its propagation velocity was found 4.7×106m/s. Therefore, the light bright emission near the wall was ascribed to excitation and ionization of air molecules due to the propagation of FIW. The reduced electrical field induced by the FIW propagation was also estimated, and it was relatively high (= 3.0×10-19V·m2) at the peak. At the same time, this high electric field duration time was found short (= 3.5ns). Furthermore, the positive polarity voltage applied to the high voltage electrode was found to provide higher reduced electric field and shorter duration time of high electric field than the negative one for gas number density less than 3.8×1024m-3. From these results, it is suggested that FIW has a possibility to become a useful ionization method for the supersonic flow.

EXPERIMENTAL AND COMPUTATIONAL INVESTIGATION OF THE AERODYNAMICS OF A SUPERSONIC PROJECTILE WITH STREAMWISE SLOTS

The paper investigates the role of streamwise slots applied to the boat-tail section of a supersonic projectile at Mach numbers from 1.36 to 1.83. The aerodynamic drag and surface pressures were measured in a supersonic indraft wind tunnel. Computations were performed at M = 1.36 and compared with the wind tunnel data.

Design methodology and performance of an indraft wind tunnel

AbstractThe design methodology and performance of Loughborough University’s new 1·9m × 1·3m, indraft wind tunnel is discussed in the following paper. To overcome severe spatial and financial constraints, a novel configuration was employed, with the inlet and exit placed adjacent to each other and opened to atmosphere. Using a fine filter mesh, honeycomb, two turbulence reduction screens and a contraction ratio of 7·3, flow uniformity in the working area of the jet at 40ms-1is shown to be within 0·3% deviation from the mean velocity, with turbulence intensity in the region of 0·15%. Working section boundary layer characteristics are shown to be consistent with that of a turbulent boundary layer growing along a flat plate, which originates at the point of inflection of the contraction. A maximum velocity of 46ms-1was achieved from a 140kW motor, compared to a prediction of 44ms-1, giving an energy ratio of 1·42. Comparison between theoretical and measured performance metrics indicate differences between the way modules perform when part of a wind tunnel system compared to data gathered from test rigs.

초음속 노즐에서 발생하는 응축충격파 진동의 피동제어

Rapid expansion of a moist air or a stream through a supersonic nozzle often leads to non-equilibrium condensation shock wave, causing a considerable energy loss in flow field. Depending on amount of latent heat released due to non-equilibrium condensation, the flow is highly unstable or a periodical oscillation accompanying the condensation shock wave in the nozzle. The unsteadiness of the condensation shock wave is always associated with several kinds of instabilities as well as noise and vibration of flow devices. In the current study, a passive control technique using a porous wall with a plenum cavity underneath is applied for the purpose of alleviation of the condensation shock oscillations in a transonic nozzle. A droplet growth equation is coupled with two-dimensional Navier-Stokes equation system. Computations are carried out using a third-order MUSCL type TVD finite-difference scheme with a second-order fractional time step. An experiment using an indraft wind tunnel is made to validate the present computational results. The results show that the oscillations of the condensation shock wave are completely suppressed by the current passive control method.

Development of the wake of an airfoil with riblets

Tests were conducted on a NACA 0012 airfoil with riblets at a freestream Reynolds number of 2.5 x 105 in an indraft wind tunnel. The riblets used in the investigation were of symmetric v-groove type with the heights of 0.0229, 0.076, and 0.152 mm. Measurements of the mean velocity profiles, turbulence, and integral quantities in the near and intermediate wakes indicated that the growth of the wake of an airfoil with riblets was similar to that of a clean airfoil. It was also observed that even at a distance of 600 momentum thicknesses downstream of the trailing edge wake equilibrium has not been reached for any of the configurations tested; however, riblet effectiveness was indicated by a marked decrease in the level of maximum shear stress, especially for the 0.152 mm riblets.

Oscillations of Laval Nozzle Flow with Condensation : Part 2, On the Mechanism of Oscillations and their Amplitudes

In the case of moist air or steam rapidly expanding in Laval nozzles and condensation occurring in the supersonic section of the flow, if the heat released by condensation exceeds a certain quantity, the flow becomes unstable and a periodic flow oscillation occurs. In the first report, an empirical relation for the frequency of the periodic flow was presented and the range of oscillations was clarified. In the present paper, the flow oscillation has been examined experimentally for the case of rapid expansion of moist air, using a Ludwieg tube and a super sonic indraft wind tunnel. As the result, the effects of the initial relative humidity of the moist air and the nozzle geometry on the amplitude of pressure oscillation have been clarified. Furthermore, based on the experimental results, the mechanism of the flow oscillation has been discussed.