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Increased turbine rotor blade loading through use of boundary layer control devices - vortex generator, tangential slot blowing, tandem airfoil, and jet flap
Understanding the unsteady aspects of turbine rotor flow fields is critical to successful future turbine designs. A technology program was conducted at NASA's Marshall Space Flight Center to increase the understanding of unsteady environments for rocket engine turbines. The experimental program involved instrumenting turbine rotor blades with miniature surface mounted high frequency response pressure transducers. The turbine model was then tested to measure the unsteady pressures on the rotor blades. The data obtained from the experimental program is unique in two respects. First, much more unsteady data was obtained (several minutes per set point) than has been possible in the past. Also, an extensive steady performance database existed for the turbine model. This allowed an evaluation of the effect of the on-blade instrumentation on the turbine's performance. A three-dimensional unsteady Navier-Stokes analysis was also used to blindly predict the unsteady flow field in the turbine at the design operating conditions and at +15 degrees relative incidence to the first-stage rotor. The predicted time-averaged and unsteady pressure distributions show good agreement with the experimental data. This unique data set, the lessons learned for acquiring this type of data, and the improvements made to the data analysis and prediction tools are contributing significantly to current Space Launch Initiative turbine airflow test and blade surface pressure prediction efforts.
A two phase flow turbine with two stages of axial flow impulse rotors was tested with three different working fluid mixtures at a shaft power of 30 kW. The turbine efficiency was 0.55 with nitrogen and water of 0.02 quality and 94 m/s velocity, 0.57 with Refrigerant 22 of 0.27 quality and 123 m/s velocity, and 0.30 with steam and water of 0.27 quality and 457 m/s velocity. The efficiencies with nitrogen and water and Refrigerant 22 were 86 percent of theoretical. At that fraction of theoretical, the efficiencies of optimized two phase turbines would be in the low 60 percent range with organic working fluids and in the mid 50 percent range with steam and water. The recommended turbine design is a two stage axial flow impulse turbine followed by a rotary separator for discharge of separate liquid and gas streams and recovery of liquid pressure.
Proposed hero turbine includes de Laval nozzles modified to new curved, longer, more-gradually-tapered shape that promotes flashing and reduces separation. Turbines designed with new nozzles compete with rotary separator turbines used in geothermal powerplants. Other potential applications include heat pumps and thermal-energy conversion systems.
Tests have been conducted on a 20-inch diameter single-stage air-cooled turbine designed to evaluate the effects of film cooling air on turbine aerodynamic performance. The present paper reports the results of five test configurations, including two different cooling designs and three combinations of cooled and solid airfoils. A comparison is made of the experimental results with a previously published analytical method of evaluating coolant injection effects on turbine performance.
An experimental investigation of a 0.8 hub-tip radius ratio, nontwisted-rotor-blade turbine designed for a stagnation-pressure ratio of 2.5 and an equivalent mean blade speed of 643 feet per second was made in a cold-air turbine with (a) nontwisted stator blades, and (b) twisted stator blades designed to maintain zero rotor-inlet incidence angles. Turbine efficiencies of the order of 0.85 at the design point were obtained with a nontwisted-rotor-blade turbine with a hub-tip radius ratio of 0.80. The turbine with the twisted stator blades gave higher efficiencies at the design point (of the order of 1.5 percentage points) than the turbine with the nontwisted stator blades.
The experimental results of the evaluation of two high work, transonic, single-stage turbines investigated under the Energy Efficient Engine (E3) Program are presented. The objective of the E3 program is to provide an advanced technology base for a new generation of fuel-conservative turbofan engines. A single-stage turbine required fewer cooled airfoils, a reduced number of leakage paths and no interstage seals. These advanced energy efficient engines require high engine pressure ratios resulting in high expansion ratio, transonic, turbine designs which must have high aerodynamic efficiency. The goal of the turbine program is to develop a high pressure turbine that is compatible with the overall engine design and has an uncooled efficiency of 90.8 percent.
The Open-Source Offshore (OSO) airfoils have been developed for research purposes for offshore wind turbines, offering a set of airfoils that align with modern turbine design requirements and industry design practices without proprietary constraints on research use. The eventual airfoil family will target the IEA 22 MW reference wind turbine, which was originally developed with the FFA airfoils. The two airfoils summarized in Table 1 (OSO-21-WT1 and OSO-30-WT1) started development as part of a family of airfoils being designed to target the IEA 22 MW wind turbine. The criteria used to design these airfoils are summarized in Table 1, which aim to encapsulate requirements of modern airfoils for offshore wind turbine applications, and were developed with feedback from industry and research experts. The airfoils were designed using XFOIL and candidate airfoils were then analyzed in RFOIL, which is considered more accurate than XFoil for high lift predictions of thicker airfoils. The design process for a preliminary family of airfoils is available, including a more detailed explanation of the design requirements and metrics similar to those used for these airfoils. Most of the design criteria are met for these two airfoils, with two exceptions. For both airfoils, the L/D Roughness Loss metric is exceeded (42% > 40% goal) and the desired lift coefficient margin over the design value (“CL_Margin”) was moderately exceeded (0.43 > 0.3) while smooth-stall characteristics (computed) were achieved. Note that all of the metrics were computed using RFOIL, and like other new airfoils, these will need to be experimentally validated at a range of Reynolds numbers. The airfoil coordinates will be shared publicly on Sadia National Laboratories’ public Github repository:
Steady state stress relaxation analysis of turbine blade cooling designs
An axial-flow turbine off-design performance computer code used for preliminary studies of gas turbine systems was modified and calibrated based on the experimental performance of large aircraft-type turbines. The flow- and loss-model modifications and calibrations are presented in this report. Comparisons are made between computed performances and experimental data for seven turbines over wide ranges of speed and pressure ratio. This report also serves as the users manual for the revised code, which is named AXOD.
A 25.4-cm (10-in) tip diameter turbine was tested to determine the effect of rotor radial tip clearance on turbine overall performance. The test turbine was a half-scale model of a 50.8-cm-(20-in.-) diameter research turbine designed for high-temperature core engine application. The test turbine was fabricated with solid vanes and blades with no provision for cooling air and tested at much reduced inlet conditions. The tests were run at design speed over a range of pressure ratios for three different rotor clearances ranging from 2.3 to 6.7 percent of the annular blade passage height. The results obtained are compared to the results obtained with three other turbines of varying amounts of reaction.
Electron beam melting (EBM) is an additive manufacturing process used to print metallic components with complex geometries for aerospace applications. A preheating characteristic enables this technology to avoid large accumulation of residual stress during printing. However, cracking mechanisms particularly in Ni-based superalloys are associated with not only residual stress but also with microstructure (i.e., solidification morphology, transient evolution of γ'). Siemens is committed to AM technology for gas turbine applications. The ability to novel predictive designs of turbine blades would reduce process development time and allow acquisition of key microstructure and performance data early in a design campaign. Siemens initiated a project to establish the design rules for the fabrication of non-printable Nibased superalloys. This phase 1 program demonstrates the feasibility for virtual fabrication of the actual blade geometries through EBM process.
The objectives of the HOST Turbine Heat Transfer subproject were to obtain a better understanding of the physics of the aerothermodynamic phenomena and to assess and improve the analytical methods used to predict the flow and heat transfer in high temperature gas turbines. At the time the HOST project was initiated, an across-the-board improvement in turbine design technology was needed. A building-block approach was utilized and the research ranged from the study of fundamental phenomena and modeling to experiments in simulated real engine environments. Experimental research accounted for approximately 75 percent of the funding with the remainder going to analytical efforts. A healthy government/industry/university partnership, with industry providing almost half of the research, was created to advance the turbine heat transfer design technology base.
NASA’s In-Space Propulsion facility experienced two infant mortality water pump bearing failures during a recent refurbishment campaign, the cause or causes of which were unclear. A forensic failure study was undertaken that included site visits, collection and examination of physical evidence and a bearing and pump design review. The pumps are large (2000 hp) vertical turbine design machines that utilize metallic sleeve bearings cooled and lubricated by the pumped fluid (water). Though initial possible root cause for the failures were attributed to improper fabrication and debris contamination, detailed examination of failure surfaces combined with bearing design revealed that the failure mode was thermal seizure brought on by inadequate cooling flow. An unusual contributing factor was that the use of grease as an assembly lubricant, as opposed to a fluid like oil, appears to have exacerbated the rather than alleviated the thermal runaway by blocking water flow. Resolution of the problem was achieved through the design and installation of dedicated bearing lubrication water feed system.
Experimental results are presented for a 0.767 scale model of the first stage of a two-stage turbine designed for a high by-pass ratio engine. The turbine was tested with both uniform inlet conditions and with an inlet radial temperature profile simulating engine conditions. The inlet temperature profile was essentially mixed-out in the rotor. There was also substantial underturning of the exit flow at the mean diameter. Both of these effects were attributed to strong secondary flows in the rotor blading. There were no significant differences in the stage performance with either inlet condition when differences in tip clearance were considered. Performance was very close to design intent in both cases.
Experimental results are presented for a 0.767 scale model of the first stage of a two-stage turbine designed for a high by-pass ratio engine. The turbine was tested with both uniform inlet conditions and with an inlet radial temperature profile simulating engine conditions. The inlet temperature profile was essentially mixed-out in the rotor. There was also substantial underturning of the exit flow at the mean diameter. Both of these effects were attributed to strong secondary flows in the rotor blading. There were no significant differences in the stage performance with either inlet condition when differences in tip clearance were considered. Performance was very close to design intent in both cases. Previously announced in STAR as N84-24589
A Users Guide for the axial flow turbine off-design computer program AXOD2 is composed in this paper. This Users Guide is supplementary to the original Users Manual of AXOD. Three notable contributions of AXOD2 to its predecessor AXOD, both in the context of the Guide or in the functionality of the code, are described and discussed in length. These are: 1) a rational representation of the mathematical principles applied, with concise descriptions of the formulas implemented in the actual coding. Their physical implications are addressed; 2) the creation and documentation of an Addendum Listing of input namelist-parameters unique to AXOD2, that differ from or are in addition to the original input-namelists given in the Manual of AXOD. Their usages are discussed; and 3) the institution of proper stoppages of the code execution, encoding termination messaging and error messages of the execution to AXOD2. These measures are to safe-guard the integrity of the code execution, such that a failure mode encountered during a case-study would not plunge the code execution into indefinite loop, or cause a blow-out of the program execution. Details on these are discussed and illustrated in this paper. Moreover, this computer program has since been reconstructed substantially. Standard FORTRAN Langue was instituted, and the code was formatted in Double Precision (REAL*8). As the result, the code is now suited for use in a local Desktop Computer Environment, is perfectly portable to any Operating System, and can be executed by any FORTRAN compiler equivalent to a FORTRAN 9095 compiler. AXOD2 will be available through NASA Glenn Research Center (GRC) Software Repository.
A gas turbine engine has an inlet duct formed to have a shape with a first ellipse in one half and a second ellipse in a second half. The second half has an upstream most end which is smaller than the first ellipse. The inlet duct has a surface defining the second ellipse which curves away from the first ellipse, such that the second ellipse is larger at an intermediate location. The second ellipse is even larger at a downstream end of the inlet duct leading into a fan.