Leading edge heat loads on turbine airfoils can be reduced by increasing the diameter of the leading edge. The lower external heat transfer and more generous curvature may allow for cooling this region internally. Large diameter leading edge regions are expected to exhibit a relatively broad region with nearly constant heat transfer. However, the ability of internal passages to cool a surface diminishes with distance as cooling air picks up thermal energy within the passage. Two novel internal cooling geometries which incrementally replenish cooling air, using impingement holes distributed along the array, have been designed and tested. These cooling methods have been compared to a baseline high solidity passage in terms of both array heat transfer and pressure drop. Heat transfer rates and pressure drop have been determined on a row by row basis to provide a means to assess their ability to sustain adequate cooling levels across the entire leading edge region. The authors believe turbine airfoil designs integrating large diameter leading edge regions with properly designed internal passages have the potential to eliminate showerhead cooling arrays in many industrial gas turbine applications. This change is especially beneficial in environments where fuel or air impurities have the potential to clog leading edge showerhead cooling arrays. Heat transfer and pressure drop measurements were acquired in a bench scale test rig. Reynolds numbers ranged from approximately 5000 to 60,000 for the constant height channel arrays based on the pin diameter and the local maximum average velocity across a row. The high solidity pin fin arrays have an axial spacing (X/D) of 1.074 and a cross channel spacing (S/D) of 1.625. The constant section pin fin arrays have channel height to diameter ratios of 0.5. Each array has eight rows of pins with six pins per row in a staggered arrangement. Heat transfer testing was conducted using a constant temperature boundary condition.
Leading edge heat loads on turbine airfoils can be reduced by increasing the diameter of the leading edge. The lower external heat transfer and more generous curvature may allow for cooling this region internally. Large diameter leading edge regions are expected to exhibit a relatively broad region with nearly constant heat transfer. However, the ability of internal passages to cool a surface diminishes with distance as cooling air picks up thermal energy within the passage. Two novel internal cooling geometries have been designed and tested, which incrementally replenish cooling air by using impingement holes distributed along the array. These cooling methods have been compared to a baseline high solidity passage in terms of both array heat transfer and pressure drop. Heat transfer rates and pressure drop have been determined on a row by row basis to provide a means to assess their ability to sustain adequate cooling levels across the entire leading edge region. The authors believe turbine airfoil designs integrating large diameter leading edge regions with properly designed internal passages have the potential to eliminate showerhead cooling arrays in many industrial gas turbine applications. This change is especially beneficial in environments where fuel or air impurities have the potential to clog leading edge showerhead cooling arrays. Heat transfer and pressure drop measurements were acquired in a bench scale test rig. Reynolds numbers ranged from approximately 5000 to 60,000 for the constant height channel arrays based on the pin diameter and the local maximum average velocity across a row. The high solidity pin fin arrays have an axial spacing (X/D) of 1.074 and a cross channel spacing (S/D) of 1.625. The constant section pin fin arrays have channel height to diameter ratios of 0.5. Each array has eight rows of pins with six pins per row in a staggered arrangement. Heat transfer testing was conducted using a constant temperature boundary condition.
Midspan heat transfer and pressure measurements have been acquired in a steady state transonic linear cascade at low to moderate Reynolds numbers. These results were used to investigate boundary layer development and separation on a 2-D variable-speed power turbine rotor blade. These distributions were acquired at 8 separate incidence angles ranging from +5.8 to -51.2. At each angle measurements were acquired at four Reynolds numbers ranging from 50,000 to 568,000 based on true chord and exit conditions. These four Reynolds number cases were tested at the design exit Mach number of 0.72. Reynolds numbers of 228,000 and 568,000 were also run for selected angles at an exit Mach number of 0.35 to constitute match points for similar measurements which were conducted at NASA Glenn. These measurements were conducted at both a lower (~0.4%) and a higher (~4%) turbulence condition. The surface pressure distributions are plotted in terms of local isentropic Mach number. They show changes in loading, the movement of the stagnation line, and regions of separation with changing incidence angle and Reynolds number. Heat transfer measurements show areas of laminar flow, regions of transition, locations of separation as well as information on reattachment. NOMENCLATUREC blade true chord length, m Cp specific heat ratio, (J/kg/K) h surface heat transfer coefficient, (W/m 2 /K) Lu energy scale, Lu = 1.5 u' 3 / Ma Mach number Re C Reynolds number based on true chord and exit conditions S surface distance from stagnation point, m St Stanton number, h/V EX Cp Tu turbulence intensity, Tu = |u'|/U IN u' rms fluctuation velocity, m/s U IN ideal cascade inlet velocity, m/s V EX ideal cascade exit velocity, m/s Greek Letter Symbols turbulent dissipation rate, m 2 /s 3 Subscripts EX refers to conditions at the nozzle exit plane IN refers to conditions at the nozzle inlet plane INTRODUCTIONThe Boeing Company estimates that worldwide air traffic will grow at approximately 5% per year over the next 20 years while North American air traffic will grow at approximately 2.7% per year over that period. Consequently, an increasing level of air traffic congestion is expected resulting in a need for new approaches in the busiest regions of the US. One approach to mitigate this congestion is with the use of a civil tilt rotor vehicle which could take off vertically and then cruise efficiently at a Mach number of approximately 0.5. NASA has published some technology goals which would help develop this possibility with its Large Civil Tilt Rotor vehicle (LCTR) [1] concept. This vehicle would take off vertically like a helicopter and later after reaching a designated altitude transition to cruise mode similar to a turboprop aircraft. One difficulty with this concept is the wide range in engine speed requirements between take-off and cruise. Over this range in the take-off and landing cycle the engine is expected to require a speed reduction as large as 50%. This would require either a variable speed transmission or a variable speed turbine. The g...
The influence of low to moderate Reynolds number and low to moderate turbulence level on aerodynamic losses is investigated in an incidence tolerant turbine blade cascade for a variable speed power turbine. This work complements midspan heat transfer and blade loading measurements which are acquired in the same cascade at the same conditions. The aerodynamic loss measurements are acquired to quantify the influence of Reynolds number and turbulence level on blade loss buckets over the wide range of incidence angles for the variable speed turbine. Eight discrete incidence angles are investigated ranging from +5.8° to −51.2°. Noting that the design inlet angle of the blade is 34.2° these incidence angles correspond to inlet angles ranging from +40° to −17°. Exit loss surveys, presented in terms of local total pressure loss and secondary velocities have been acquired at four exit chord Reynolds numbers ranging from 50,000 to 568,000. These measurements were acquired at both low (∼0.4%) and moderate (∼4.0%) inlet turbulence intensities. The total pressure losses are also presented in terms of cross passage averaged loss and turning angle. The resulting loss buckets for passage averaged losses are plotted at varied Reynolds numbers and turbulence condition. The exit loss data quantify the impact of Reynolds number and incidence angle on aerodynamic losses. Generally, these data document the substantial deterioration of performance with decreasing Reynolds number.
This research is focused on the effects of large changes in Reynolds number that typically occurs during the flight of high altitude UAV's. This paper documents the influence of Reynolds number, turbulence level, and exit Mach number on the vane surface Stanton number. Reynolds number is based on true chord and exit conditions and ranges from 90,000 through 1,000,000. Low and high inlet turbulence levels were developed for the study and determined to be 0.8% and 9.0%. Tests were run at exit Mach numbers of 0.7, 0.8 and 0.9. These surface heat transfer measurements were acquired in the University of North Dakota's transonic cascade test facility. This facility uses a closed loop to allow the regulation of system pressure to control the test condition Reynolds number. The Mach number is adjusted using a "roots" blower driven by a variable frequency drive. Heat transfer measurements were acquired using a constant heat flux foil fabricated using a 0.023 mm Inconel foil backed with 0.05 mm of Kapton and adhered to the heat transfer vane using a high temperature acrylic adhesive. The linear cascade is configured in a four vane three full passage arrangement. The low turbulence condition is developed using the existing flow conditioning section coupled to a 4.7 to 1 area ratio nozzle. The high turbulence condition uses a mock aero combustor to generate a turbulence level of around 9.0%. These data show the influence of Mach number, Reynolds number and turbulence level on transition and heat transfer augmentation and are expected to be useful in grounding heat transfer predictive methods applicable to small or high altitude gas turbine engines. NomenclatureC vane true chord length, m Cp specific heat ratio, (J/kg/K) h surface heat transfer coefficient, (W/m 2 /K) Lu energy scale, Lu = 1.5 u' 3 / Ma Mach number Re C Reynolds number based on true chord and exit conditions S surface distance from stagnation point, m St Stanton number, h/V EX Cp Tu turbulence intensity, Tu = |u'|/U IN u' rms fluctuation velocity, m/s U IN ideal cascade inlet velocity, m/s V EX ideal cascade exit velocity, m/s Y Normal distance from the wall, m Y + normal distance in wall units, Y + = Y (/)/ Greek Letter Symbols turbulent dissipation rate, m 2 /s 3 kinematic viscosity, m 2 /s density, kg/m 3 surface shear stress, N/m 2 Subscripts EX refers to conditions at the nozzle exit plane IN refers to conditions at the nozzle inlet plane INTRODUCTIONUnmanned aerial vehicles (UAVs) have a growing mission in supporting military engagements abroad as well as ensuring domestic security in the US. UAV's have been engaged to help assess the damage of natural disasters, patrol borders and assess crop health. The yearly market for unmanned aerial vehicles (UAV's) is expected to grow to at a rate of 12% through 2018 to a value of $18.6 billion according to Market Research Media. A growing number of large to medium UAV's are being designed to use turbofan engines due to the ability to achieve high thrust to weight ratios and high thermal and propu...
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