Nirm Velumylum Nirmalan

The Measurement of Full-Surface Internal Heat Transfer Coefficients for Turbine Airfoils Using a Non-Destructive Thermal Inertia Technique

A new method has been developed and demonstrated for the non-destructive, quantitative assessment of internal heat transfer coefficient distributions of cooled metallic turbine airfoils. The technique employs the acquisition of full-surface external surface temperature data in response to a thermal transient induced by internal heating/cooling, in conjunction with knowledge of the part wall thickness and geometry, material properties, and internal fluid temperatures. An imaging Infrared camera system is used to record the complete time history of the external surface temperature response during a transient initiated by the introduction of a convecting fluid through the cooling circuit of the part. The transient data obtained is combined with the cooling fluid network model to provide the boundary conditions for a finite element model representing the complete part geometry. A simple 1D lumped thermal capacitance model for each local wall position is used to provide a first estimate of the internal surface heat transfer coefficient distribution. A 3D inverse transient conduction model of the part is then executed with updated internal heat transfer coefficients until convergence is reached with the experimentally measured external wall temperatures as a function of time. This new technique makes possible the accurate quantification of full-surface internal heat transfer coefficient distributions for prototype and production metallic airfoils in a totally non-destructive and non-intrusive manner. The technique is equally applicable to other material types and other cooled/heated components.Copyright

Multicolor Techniques for Identification and Filtering of Burst Signals in Jet Engine Pyrometers

A Defense Advanced Research Projects Agency (DARPA)-funded multicolor pyrometry (MCP) experiment was carried out on a government-provided aircraft engine to study the nature of hot particulate bursts generated from the combustor at certain engine conditions. These bursts of hot particulates lead to intermittent high-voltage signal output from the line-of-sight (LOS) pyrometer that is ultimately detected and used by the onboard digital engine controller (DEC). The investigation used a high-speed MCP system designed to detect bursts and identify their properties. Results of the radiant temperature, multicolor temperature, and apparent emissivity are presented. The results indicated that the apparent emissivity calculated during the signal burst was lower than that of the blade. The root cause for the signal burst was identified as soot particles generated as a by-product of combustion under certain conditions. This conclusion was drawn based on both experimental and simulation results. Technical strategies to separate, reduce, or remove the burst signal are proposed.

The Measurement of Full-Surface Internal Heat Transfer Coefficients for Turbine Airfoils Using a Nondestructive Thermal Inertia Technique

A new method has been developed and demonstrated for the non-destructive, quantitative assessment of internal heat transfer coefficient distributions of cooled metallic turbine airfoils. The technique employs the acquisition of full-surface external surface temperature data in response to a thermal transient induced by internal heating/cooling, in conjunction with knowledge of the part wall thickness and geometry, material properties, and internal fluid temperatures. An imaging Infrared camera system is used to record the complete time history of the external surface temperature response during a transient initiated by the introduction of a connecting fluid through the cooling circuit of the part. The transient data obtained is combined with the cooling fluid network model to provide the boundary conditions for a finite element model representing the complete part geometry. A simple 1-D lumped thermal capacitance model for each local wall position is used to provide a first estimate of the internal surface heat transfer coefficient distribution. A 3-D inverse transient conduction model of the part is then executed with updated internal heat transfer coefficients until convergence is reached with the experimentally measured external wall temperatures as a function of time. This new technique makes possible the accurate quantification of full-surface internal heat transfer coefficient distributions for prototype and production metallic airfoils in a totally nondestructive and non-intrusive manner The technique is equally applicable to other material types and other cooled/heated components.

Lower heating value sensor for fuel monitoring

This work details the development of a low cost, lower heating value (LHV) sensor that consists of a catalytic film deposited on the surface of a micromachined hotplate. The micromachined sensor has low heat capacity and thermal conductivity, making it appropriate for accurate LHV determination. Catalytically combusting the fuel provides the capability of direct LHV measurement, unlike most LHV measurement systems which rely on inference. The results of extensive laboratory testing and preliminary field testing will be reported, which demonstrate the capability to measure LHV values to within +/- 5% accuracy across a wide range of fuel values. Field testing results have shown LHV determination to within +/- 1.2% over a more narrow range of fuel heating values

Multi-Color Techniques for Identification and Filtering of Burst Signals in Jet Engine Pyrometers

A DARPA funded Multi-color Pyrometry (MCP) experiment was carried out on a government provided aircraft engine to study the nature of hot particulate bursts generated from the combustor at certain engine conditions. These bursts of hot particulates lead to intermittent high-voltage signal output from the line-of-sight (LOS) pyrometer which is ultimately detected and used by the onboard digital engine controller (DEC). The investigation used a high-speed MCP system designed to detect bursts and identify their properties. Results of the radiant temperature, multi-color temperature and apparent emissivity are presented. The results indicated that the apparent emissivity calculated during the signal burst was lower than that of the blade. The root cause for the signal burst was identified as soot particles generated as by-product of combustion under certain conditions. This conclusion was drawn based on both experimental and simulation results. Technical strategies to separate, reduce or remove the burst signal are proposed.Copyright

Passive Absorption/Emission Spectroscopy for Gas Temperature Measurements in Gas Turbine Engines

A novel technique is developed to simultaneously measure hot surface and gas temperatures based on passive absorption/emission spectroscopy (PAS). This non-intrusive, in situ technique is the extension of multi-wavelength pyrometry to also measure gas temperature. The PAS technique uses hot surface (e.g., turbine blade) as the radiation source, and measures radiation signals at multiple wavelengths. Radiation signals at wavelengths with minimum interference from gas (mostly from water vapor and CO2 ) can be used to determine the hot surface temperature, while signals at wavelengths with gas absorption/emission can be used to determine the gas temperature in the line-of-sight. The detection wavelengths are optimized for accuracy and sensitivity for gas temperature measurements. Simulation results also show the effect of non-uniform gas temperature profile on measurement results. High pressure/temperature tests are conducted in single nozzle combustor rig to demonstrate sensor proof-of-concept. Preliminary engine measurement results shows the potential of this measurement technique. The PAS technique only requires one optical port, e.g., existing pyrometer or borescope port, to collect the emission signal, and thus provide practical solution for gas temperature measurement in gas turbine engines.Copyright

Experimental and Numerical Study of Heat Transfer in a Gas Turbine Combustor Liner

Experiments and numerical simulations were conducted to understand the heat transfer characteristics of a stationary gas turbine combustor liner cooled by impingement jets and cross flow between the liner and sleeve. Heat transfer was also aided by trip-strip turbulators on the outside of the liner and in the flowsleeve downstream of the jets. The study was aimed at enhancing heat transfer and prolonging the life of the combustor liner components. The combustor liner and flow sleeve were simulated using a flat plate rig. The geometry has been scaled from actual combustion geometry except for the curvature. The jet Reynolds number and the mass-velocity ratios between the jet and cross flow in the rig were matched with the corresponding combustor conditions. A steady state liquid crystal technique was used to measure spatially resolved heat transfer coefficients for the geometric and flow conditions mentioned above. The heat transfer was measured both in the impingement region as well as over the turbulators. A numerical model of the combustor test rig was created that included the impingement holes and the turbulators. Using CFD, the flow distribution within the flow sleeve and the heat transfer coefficients on the liner were both predicted. Calculations were made by varying the turbulence models, numerical schemes, and the geometrical mesh. The results obtained were compared to the experimental data and recommendations have been made with regard to the best modeling approach for such liner-flow sleeve configurations.Copyright

Experimental Investigation of Aerodynamic Losses of Different Shapes of a Shrouded Blade Tip Section

An experimental investigation was conducted to study the effects on aerodynamic losses of different tip shroud shapes of a shrouded turbine blade. Pressures were measured on the airfoil surface near the tip and a plane downstream of the exit plane in a three-airfoil stationary cascade. The instrumented center airfoil and the two slave airfoils modeled the aerodynamic tip section of a blade and have the capability to vary tip clearance. The experiments were run at tip-clearances varying from 0.25% to 1.67% and at an exit Reynolds number of 1.25 × 106 and Mach Number of 0.95. The paper presents the influence of three tip-shroud shapes and five different tip-clearances on the aerodynamic losses.Copyright

Experimental and Computational Investigation of Heat Transfer Effectiveness and Pressure Distribution of a Shrouded Blade Tip Section

An experimental and computational investigation was conducted to study the detailed distribution of heat transfer effectiveness and pressure on an attached tip-shroud of a turbine blade. Temperatures and pressures were measured on the airfoil-side and gap-side surfaces of the shrouded tip in a three-airfoil stationary cascade. The instrumented center airfoil and the two slave airfoils modeled the aerodynamic tip section of a blade and have the capability to vary tip clearance. The experiments were run at gaps varying of 0.25% to 1.67% of blade span and at an airfoil exit Reynolds number of 1.26×106 and Mach number of 0.95. The effect of coolant flow through the radial-cooled airfoil was also studied. The experimental results are compared with a computational model using the commercially available code, CFX. This unique study presents the influence of gap and coolant flow on the pressure distribution and heat transfer effectiveness of an attached tip-shroud surface.Copyright

The Determination of In-Situ Film Hole Flow Rates Using a Transient Thermal Inertia Method

Film hole flow rates are conventionally characterized by a discharge coefficient relating the actual mass flow rate to the theoretical ideal flow rate based upon some measured effective hole diameter or flow area. These discharge coefficients are typically measured on controlled test plates that contain the particular size, shape, and fabrication method for an individual film hole type. Such discharge coefficients are then assumed to apply to all of the in-situ film holes of that type which are machined or formed in the as-fabricated cooled turbine component. The thermal-mechanical analysis of the component is then performed using these assumed values to calculate film hole flow rates. In practice however, every film hole in a cooled airfoil is different due to machine tool wear, surface curvatures, laser drift, coating variations, and local flow supply behavior. A new method has been developed and demonstrated which allows determination of the individual film hole flow rates in-situ for an as-fabricated component, thus avoiding the need for assumed discharge coefficients or highly detailed flow checks. This method uses the thermal transient characteristics of external surface points near an active film hole to determine the flow rate through the hole. An imaging Infrared system is used to record the component response to an induced thermal cooling transient in which the film hole internal heat transfer dominates the local thermal transient behavior. The characteristic of the non-dimensionalized thermal decay is related to the flow rate within each individual film hole using a limited calibration function. This method allows the rapid inspection and quantification of detailed film hole flows for actual parts, which data may then be used in the analysis and health monitoring of parts in operation.Copyright