Wind Tunnel Free Stream Research Articles

Aerodynamic characteristics of the retro propulsion landing burn of vertically landing launchers

In the frame of the European funded H2020 project RETALT (retro propulsion-assisted landing technologies), the unsteady aerodynamics of vertically descending and landing launchers have been investigated. In this paper, experimental data of the landing burn tested in the Vertical Free-Jet Facility Cologne at DLR in Cologne are presented. The landing burn was simulated with a cold gas jet of pressurized air opposing the wind tunnel free stream. Tests with several jet conditions were compared to results without active jet. Proper orthogonal decomposition of schlieren recordings and spectral analyses of their time histories are performed and are compared to frequencies in pressure measurements. Dominant frequencies were found, which are strongest at Mach 0.8. Especially, a Strouhal number of 0.2 was found to be most dominant. The intensity of the dominant frequencies can be lowered if the engine is active. The normalized root mean square pressure fluctuations are between 0.1 and 0.3 during the landing maneuver. Additionally, the steady flow features scale well with the ambient pressure ratio and the momentum flux ratio. The unsteady flow field dynamics of the subsonic retro propulsion flow field can likely be linked to large-scale turbulent structures in the supersonic jet, triggering large-scale pressure fluctuations and altering the overall flow field.

Spatially-Resolved Freestream Velocity Measurements At The NASA Langley 31-Inch Mach 10 Air Tunnel Using FLEET

Freestream velocity measurements in the NASA Langley 31-inch Mach 10 wind tunnel are reported in this paper using Femtosecond Laser Electronic Excitation Tagging (FLEET). The freestream measurements acquired during the January 2023 test campaign were the first direct measurement of freestream velocity in this hypersonic wind tunnel facility. Spatial distributions of time-averaged and instantaneous velocity measurements were obtained at all three typical wind tunnel freestream unit Reynolds number conditions of Re∞ /L = 1800000 [1/m] , 3600000 [1/m], and 6400000 [1/m], though the current paper focuses on centerline measurements for the three Re∞ /L and spatial distributions for one Re∞ /L. Measured values for time-averaged velocity and mean of the instantaneous velocity at the wind tunnel centerline agree within 5 m/s or 0.4% of the calculated velocity from the facility data acquisition system. Measurements acquired at locations away from the wind tunnel centerline reveal the spatial extent of the core flow of the hypersonic facility.

Effect of turbulence on the performance of VAWTs: An experimental study in two different wind tunnels

Turbulence is one of the main characteristics of urban flows, but its effect on the performance of small VAWTs (often used in urban wind installations) has not been thoroughly researched. This experimental study focuses in testing a H-Darrieus VAWT prototype in different turbulent conditions inside the wind tunnel, in order to study the influence of turbulence intensity, integral length scales and Reynolds number. Passive grids are used to increase the wind tunnel free stream turbulence up to Iu = 15%, with integral length scales of Lux = 0.18 m and Reynolds numbers of ReD ≈ 300,000. The measurements are repeated in two wind tunnels of different size, which strengthens the results and helps quantify the effect of blockage on the turbine ratings. The results show power coefficient increases up to 20% from smooth (Iu = 0.5%) to turbulent (Iu = 15%) flows, an effect that is enhanced at low λ but that fades as ReD > 400,000. The study of the crossed influence of ReD, Iu and Lux offers valuable data in the process of optimizing the operation of small VAWTs inside urban environments.

Experimental investigation on aero-heating of rudder shaft within laminar/turbulent hypersonic boundary layers

The aero-heating of the rudder shaft region of a hypersonic vehicle is very harsh, as the peak heat flux in this region can be even higher than that at the stagnation point. Therefore, studying the aero-heating of the rudder shaft is of great significance for designing the thermal protection system of the hypersonic vehicle. In the wind tunnel test of the aero-heating effect, we find that with the increase of the angle of attack of the lifting body model, the increasement of the heat flux of the rudder shaft is larger under laminar flow conditions than that under turbulent flow conditions. To understand this, we design a wind tunnel experiment to study the effect of laminar/turbulent hypersonic boundary layers on the heat flux of the rudder shaft under the same wind tunnel freestream conditions. The experiment is carried out in the ∅2 m shock tunnel (FD-14A) affiliated to the China Aerodynamics Research and Development Center (CARDC). The laminar boundary layer on the model is triggered to a turbulent one by using vortex generators, which are 2 mm-high diamonds. The aero-heating of the rudder shaft (with the rudder) and the protuberance (without the rudder) are studied in both hypersonic laminar and turbulent boundary layers under the same freestream condition. The nominal Mach numbers are 10 and 12, and the unit Reynolds numbers are 2.4 × 106 m−1 and 2.1 × 106 m−1. The angle of attack of the model is 20°, and the deflection angle of the rudder and the protuberance is 10°. The heat flux on the model surface is measured by thin film heat flux sensors, and the heat flux distribution along the center line of the lifting body model suggests that forced transition is achieved in the upstream of the rudder. The test results of the rudder shaft and the protuberance show that the heat flux of the rudder shaft is lower in the turbulent flow than that in the laminar flow, but the heat flux of the protuberance is the other way around, i.e., lower in the laminar flow than in the turbulent flow. The wind tunnel test results is also validated by numerical simulations. Our analysis suggests that this phenomenon is due to the difference of boundary layer velocities caused by different thickness of boundary layer between laminar and turbulent flows, as well as the restricted flow within the rudder gap. When the turbulent boundary layer is more than three times thicker than that of the laminar boundary layer, the heat flux of the rudder shaft under the laminar flow condition is higher than that under the turbulent flow condition. Discovery of this phenomenon has great importance for guiding the design of the thermal protection system for the rudder shaft of hypersonic vehicles.

Supersaturation Total Temperature, Pitot Pressure, and Rayleigh Scattering Measurements at Mach 10

A fundamental research study of the supersaturation region is presented. This region is mapped in the NASA Langley 31-Inch Mach 10 Air Wind Tunnel freestream flow using physical probes (total temperature and pitot) and a nonintrusive offbody diagnostic (laser Rayleigh scattering). Data from all three methods are acquired simultaneously. Facility stagnation pressures spanned 1.0–10.0 MPa (150–1450 psi), and stagnation temperatures spanned 357–1000 K (184–1350°F). Each instrument has its own unique supersaturation region over which quantitative measurements can be obtained. The extent of each supersaturation region is a unique function of the sensitivity of the instrument selected to flowfield clustering. Laser-Rayleigh-scattering-measured density results agree to better than 15% with the freestream density computed using the GASPROPS code over a significant fraction of the supersaturation region. Evidence is presented showing that the “frozen” vibrational nonequilibrium Boltzmann population of and molecules inhibit clustering and condensation. This is the physical mechanism responsible for creating the supersaturation region. In direct conflict with the assumption in all hypersonic literature from the past 50 years, evidence is presented indicating that no hypersonic or hypervelocity freestream in which nucleation is occurring contains a frozen vibrational nonequilibrium population. This population is effectively eliminated by interacting with clusters.

Influence of wind tunnel freestream turbulence level on boundary-layer separation

An investigation of aerodynamic performance for micro air vehicle (MAV) wing is performed. It is shown, what is a reason of significant wing performance difference between tests, which were carried out in different wind tunnels. New method for estimation of wind tunnel condition influence on experimental results is suggested.

A108 複合感圧塗料計測システムの研究開発

Bi-luminophore Pressure-Sensitive Paint (bi-PSP) has been developed in order to correct the error due to temperature dependence of PSP and increase PSP measurement accuracy. The bi-PSP was composed of pressure- and temperature-sensitive dye. Tetranuclear europium (III) complexes and PdTFPP were used as temperature- and pressure- sensitive dye, respectively. The luminescence intensity of the Eu complex was highly sensitive to temperature and insensitive to pressure. The bi-PSP was examined using a painted coupon and its characteristics were clarified. As a verification test, pressure distributions on a supersonic transport (SST) model at low-speed flow were acquired by the bi-PSP measurement system. The root mean square of Cp at 50m/s was approximately 0.1 (150 Pa) at flow speed of 50m/s, indicating that this bi-PSP system can measure more accurately pressure than conventional pressure conversion technique using reference PSP images immediately after stop of wind tunnel free stream.

Experimental research on blunt trailing-edge airfoil sections at low Reynolds numbers

The wind-tunnel test results of five airfoil sections with 28.57% chord thickness at chord Reynolds number of 33,000, 66,000, and 99,000 are presented. Lift and drag forces have been measured and flow visualizations have been made. It is shown by cutting off the trailing edge of the section and by making its trailing edge blunt at low Reynolds numbers that the total drag can be reduced, the maximum lift increased, the linearity of lift curve with incidences improved, and the maximum lift-to-drag ratio increased. The structural merits of blunt trailing-edge sections are also discussed. Nomenclature C = chord length, 84 mm CD = drag coefficient CL = lift coefficient Cvir = chord length of the original section Re = chord Reynolds number, UC/v t = maximum thickness, 24 mm U = wind-tunnel freestream velocity ot = angle of attack, deg

Wind tunnel simulation of airflow and pollutant diffusion over complex terrain

A wind tunnel simulation of airflow and pollutant diffusion has been conducted for Mt. Tsukuba, a rather isolated mountain. the wind tunnel simulation was carried out in the Thermally Stratified Wind Tunnel of the National Research Institute for Pollution and Resources. The scale of the topographic model was 1/7500 for the horizontal and 1/5000 for the vertical. A nocturnal inversion was simulated in the wind tunnel by heating the airflow to 40°C and chilling the floor of the test section at 10°C. The wind tunnel free stream speed was 0.5 m s −1. The airflow was found to be satisfactorily simulated in the wind tunnel. The upstream flow stagnation, the dividing streamline and the wake in the lee of the mountain were reproduced.

Review of wind-tunnel freestream pressure fluctuations

(lib) For this flow condition, a saturated laser, K}^ , andpw>5 Tor r cm, i.e., f//2>5, the axisymmetric nozzle has an output power twice that of a two-dimensional nozzle with the same exit flow conditions and exit width w. In this pw regime, however, the chemical efficiency of the axisymmetric nozzle is 20% or less that of a premixed (pw = 0) laser. For pw = 1 Torr cm, i.e., fl/2 = l, the axisymmetric nozzle has about 30% greater output than the corresponding two-dimensional nozzle and a chemical efficiency equal to about 60% that of a premixed laser. For /?w /2<0.5, both nozzles have essentially premixed laser performance. For either nozzle, a reduction in pw results in improved performance (Fig. 3). The effect of p/pF*\0 is found in Eq. (Ha). Fabrication and nozzle wall boundary-layer effects also have to be considered when comparing two-dimensional and axisymmetric nozzles. These are beyond the scope of this study. Acknowledgment

Hypersonic Transitional Boundary Layers

Surface thin-film gages have been used to determine the extent of the transition region, intermittency distribution, and disturbance convection velocities in the boundary layer on a sharp 5-deg half angle cone at Mach 7.4 in the Ames 3.5-ft Hypersonic Wind Tunnel. In addition, extensive hot wire space-time correlation measurements have been obtained in the wind-tunnel freestream and in the transitional hypersonic boundary layer on a cone-ogive-cylinder in the same facility. Disturbance convection velocities have been obtained from the longitudinal cross correlation measurements as a function of fluctuation scale and distance from the wall. The results of normal cross correlation measurements are also discussed.