Experimental investigation on static/dynamic characteristics of a fast-response pressure sensitive paint

An optical-based technique using Pressure-Sensitive Paint（PSP） is a promising method to measure the distribution of surface pressure on an aerodynamic model. The static and dynamic characteristics of a fast-response PSP that is developed in the Chinese Academy of Sciences（CAS）are analyzed and tested to serve as the basis for experiments on unsteady surface measurement using a fast-response PSP. Two calibration systems used for this study are set up to investigate the temperature dependency, response time, and resolution. A data processing method, used for dynamic data, is analyzed and selected carefully to determine the optimum signal. Results show that the fastresponse PSP can be used normally at temperatures from 25 ℃ to 80 ℃. The effect of temperature on the accuracy of the measurement must be considered when temperatures are beyond the temperature range of 30–40 ℃. The dynamic calibration device with a solenoid valve can achieve a pressure jump within a millisecond order. The resolution is determined by the signal-to-noise ratio of the photo-multiplier tube. Results of the measurement show that the response time of the PSP decreases with a large pressure variation, and the response time is below 0.016 s when the pressure variation is under 40 kPa.

Experimental Investigation of a Linear Cascade with Large Solidity Using Pressure Sensitive Paint and Dual-Camera System

Pressure Sensitive Paint(PSP)technique has been increasingly applied to the experimental research of aerodynamics and thermodynamics due to its strengths of non-contact,high resolution results and large coverage area,etc.However,rarely has this technique been successfully used to the study of internal flow such as compressor cascade,since narrow flow passages would heavily restrict the acquisition of PSP images.In this paper,PSP technique was used to study the pressure distribution on a linear compressor cascade with large solidity of 2.3,where the view of recording camera can be heavily blocked due to adjacent blade surfaces.To help get integrated PSP images of the internal flow passage,dual camera system along with image processing tools like 3D reconstruction and image integration were adopted.The results showed that with the aid of such assistance,image results with good quality and readability could be obtained.Meanwhile,pressure data given by PSP were compared with data from traditional way of pressure taps and showed good consistency.Massive results of the entire cascade passage surface were given with different inlet Mach numbers and incidence angles.The results showed that PSP technique can integrally measure cascade tunnel of large solidity with the help of dual-camera system.

Recent Developments of Image Based Measurement Methods for Application to Transonic Flows in Industrial Wind Tunnels

The experimental investigation of unsteady complex flow fields in wind tunnels requires advanced measurement techniques. The most important of such image based measurement techniques are those for the measurement of planar flow velocity fields, planar pressure distribution, model location and deformation, model temperature and quantitative high speed flow visualization. The applications as carried out by DLR range from low speed flows to transonic flows, from high lift configurations to propellers and rotors, from wake vortex investigations in catapult facilities and water towing tanks to investigations of vortex break down phenomena on delta wings. The capability to use image based measurement techniques in transonic flows requires dedicated technical developments and experienced scientists due to the special environment of a transonic wind tunnel. In this paper an overview of the state-of-the art of the application of image based measurement techniques in transonic flows as performed by DLR＇s Institute of Aerodynamics and Flow Technology will be given.

Film cooling adiabatic effectiveness measurements of pressure side trailing edge cooling configurations

Nowadays total inlet temperature of gas turbine is far above the permissible metal temperature;as a consequence,advanced cooling techniques must be applied to protect from thermal stresses,oxidation and corrosion the components located in the high pressure stages,such as the blade trailing edge.A suitable design of the cooling system for the trailing edge has to cope with geometric constraints and aerodynamic demands;state-of-the-art of cooling concepts often use film cooling on blade pressure side:the air taken from last compressor stages is ejected through discrete holes or slots to provide a cold layer between hot mainstream and the blade surface.With the goal of ensuring a satisfactory lifetime of blades,the design of efficient trailing edge film cooling schemes and,moreover,the possibility to check carefully their behavior,are hence necessary to guarantee an appropriate metal temperature distribution.For this purpose an experimental survey was carried out to investigate the film covering performance of different pressure side trailing edge cooling systems for turbine blades.The experimental test section consists of a scaled-up trailing edge model installed in an open loop suction type test rig.Measurements of adiabatic effectiveness distributions were carried out on three trailing edge cooling system configurations.The baseline geometry is composed by inclined slots separated by elongated pedestals;the second geometry shares the same cutback configuration,with an additional row of circular film cooling holes located upstream;the third model is equipped with three rows of in-line film cooling holes.Experiments have been performed at nearly ambient conditions imposing several blowing ratio values and using carbon dioxide as coolant in order to reproduce a density ratio close to the engine conditions(DR¼1.52).To extend the validity of the survey a comparison between adiabatic effectiveness measurements and a prediction by correlative approach was performed to compare the experimental results with 1D methodologies.

Reconstruction of skin friction topology in complex separated flows

This paper describes a theoretical method for reconstruction of the skin friction topology in complex separated flows,which is developed based on the exact relation between skin friction and surface pressure through the boundary enstrophy flux(BEF).The key of this method is that a skin friction field is reconstructed from a surface pressure field as an inverse problem by applying a variational method.For applications,the approximate method is proposed,where the composite surface pressure field is given by a linear superposition of the base-flow surface pressure field and the surface pressure variation field and the base-flow BEF field is used as the first-order approximation.This approximate method is constructive in a mathematical sense since a complex skin friction field in separated flows can be reconstructed from some elemental skin friction structures(skin friction source/sink,vortex and their combinations)by a linear superposition of some simple surface pressure structures.The distinct topological features,such as critical points,separation lines and attachment lines,naturally occur as a result of such reconstruction.As examples,some elemental skin friction structures in separated flows are reconstructed in simulations,and the skin friction fields in shock-wave/boundary-layer interactions(SWBLIs)are reconstructed from pressure sensitive paint(PSP)images obtained in wind tunnel experiments.

An experimental investigation on the trailing edge cooling of turbine blades

An experimental study was conducted to quantify the flow characteristics of the wall jets pertinent to trailing edge cooling of turbine blades.A high-resolution stereoscopic particle image velocimetry(PIV)system was used to conduct detailed flow field measurements to quantitatively visualize the evolution of the unsteady vortices and turbulent flow structures in the cooling wall jet streams and to quantify the dynamic mixing process between the cooling jet stream and the mainstream flows.The detailed flow field measurements were correlated with the adiabatic cooling effectiveness maps measured by using pressure sensitive paint(PSP)technique to elucidate underlying physics in order to explore/optimize design paradigms for improved cooling effectiveness to protect the critical portions of turbine blades from harsh environments.

Turbine Endwall Film Cooling With Combustor-Turbine Interface Gap Leak- age Flow: Effect of Incidence Angle

This paper is focused on the film cooling performance of combustor-turbine leakage flow at off-design condition. The influence of incidence angle on film cooling effectiveness on first-stage vane endwall with combustor-turbine interface slot is studied. A baseline slot configuration is tested in a low speed four-blade cascade comprising a large-scale model of the GE-E 3 Nozzle Guide Vane (NGV). The slot has a forward expansion angle of 30 deg. to the endwall surface. The Reynolds number based on the axial chord and inlet velocity of the free-stream flow is 3.5 × 10 5 and the testing is done in a four-blade cascade with low Mach number condition (0.1 at the inlet). The blowing ratio of the coolant through the interface gap varies from M = 0.1 to M = 0.3, while the blowing ratio varies from M = 0.7 to M = 1.3 for the endwall film cooling holes. The film-cooling effectiveness distributions are obtained using the pressure sensitive paint (PSP) technique. The results show that with an increasing blowing ratio the film-cooling effectiveness increases on the endwall. As the incidence angle varies from i = +10 deg. to i = 10 deg., at low blowing ratio, the averaged film-cooling effectiveness changes slightly near the leading edge suction side area. The case of i = +10 deg. has better film-cooling performance at the downstream part of this region where the axial chord is between 0.15 and 0.25. However, the disadvantage of positive incidence appears when the blowing ratio increases, especially at the upstream part of near suction side region where the axial chord is between 0 and 0.15. On the main passage endwall surface, as the incidence angle changes from i = +10 deg. to i = 10 deg., the averaged film-cooling effectiveness changes slightly and the negative incidence appears to be more effective for the downstream part film cooling of the endwall surface where the axial chord is between 0.6 and 0.8.