Accession Number : ADA504905


Title :   Pulsed Film Cooling on a Turbine Blade Leading Edge


Descriptive Note : Doctoral thesis


Corporate Author : AIR FORCE INST OF TECH WRIGHT-PATTERSON AFB OH GRADUATE SCHOOL OF ENGINEERING AND MANAGEMENT


Personal Author(s) : Rutledge, James L


Full Text : https://apps.dtic.mil/dtic/tr/fulltext/u2/a504905.pdf


Report Date : Sep 2009


Pagination or Media Count : 275


Abstract : Unsteadiness in gas turbine film cooling jets may arise due to inherent unsteadiness of the flow through an engine or may be induced as a means of flow control. The traditional technique used to evaluate the performance of a steady film cooling scheme is demonstrated to be insufficient for use with unsteady film cooling and is modified to account for the cross coupling of the time dependent adiabatic effectiveness and heat transfer coefficient. The addition of a single term to the traditional steady form of the net heat flux reduction equation with time averaged quantities accounts for the unsteady effects. An experimental technique to account for the influence of the new term was devised and used to measure the influence of a pulsating jet on the net heat flux in the leading edge region of a turbine blade. High spatial resolution data was acquired in the near-hole region using infrared thermography coupled with experimental techniques that allowed application of the appropriate thermal boundary conditions immediately adjacent to the film cooling hole. The turbine blade leading edge was simulated by a half cylinder in cross flow with a blunt afterbody. The film cooling geometry consisted of a coolant hole located 21.5? from the leading edge, angled 20? to the surface and 90? from the streamwise direction. Investigated parameters include pulsation frequency, duty cycle, and waveform shape. Separate experiments were conducted in a water channel to provide visualization of the unsteady coolant propagation behavior. Further insight into the flow physics was obtained through computational simulations of the experimental apparatus. The computational results afforded time resolved flow field and net heat flux reduction data unobtainable with the experimental techniques. A technique to predict the performance of an unsteady film cooling scheme through knowledge of only the steady film cooling behavior was developed and demonstrated to be effective.


Descriptors :   *TURBINE BLADES , *LEADING EDGES , *FILM COOLING , THESES , GAS TURBINES , FLOW FIELDS , INFRARED RADIATION , HEAT FLUX , THERMAL BOUNDARY LAYER , THERMOGRAPHY , HEAT TRANSFER COEFFICIENTS , WAVEFORMS , CROSS FLOW , HEAT TRANSFER , TIME DEPENDENCE


Subject Categories : Aircraft
      Hydraulic and Pneumatic Equipment


Distribution Statement : APPROVED FOR PUBLIC RELEASE