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  • 1
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    Results of an Advanced Fan Stage Operating Over a Wide Range of Speed and Bypass Ratio (2011)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: NASA s Fundamental Aeronautics Program is investigating turbine-based combined cycle (TBCC) propulsion systems for access to space because it provides the potential for aircraft-like, space-launch operations that may significantly reduce launch costs and improve safety. To this end, National Aeronautics and Space Administration (NASA) and General Electric (GE) teamed to design a Mach 4 variable cycle turbofan/ramjet engine for access to space. To enable the wide operating range of a Mach 4+ variable cycle turbofan ramjet required the development of a unique fan stage design capable of multi-point operation to accommodate variations in bypass ratio (10 ), fan speed (7 ), inlet mass flow (3.5 ), inlet pressure (8 ), and inlet temperature (3 ). In this paper, NASA has set out to characterize a TBCC engine fan stage aerodynamic performance and stability limits over a wide operating range including power-on and hypersonic-unique "windmill" operation. Herein, we will present the fan stage design, and the experimental test results of the fan stage operating from 15 to 100 percent corrected design speed. Whereas, in the companion paper, we will provide an assessment of NASA s APNASA code s ability to predict the fan stage performance and operability over a wide range of speed and bypass ratio.
    Keywords: Aerodynamics
    Type: NASA/TM-2011-216769/PART1 , GT2010-22825 , E-17395-1 , Turbo Expo 2010; Jun 14, 2010 - Jun 18, 2010; Glasgow, Scotland; United States
    Format: application/pdf
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  • 2
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    Active Flow Separation Control of a Stator Vane Using Surface Injection in a Multistage Compressor Experiment (2003)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Micro-flow control actuation embedded in a stator vane was used to successfully control separation and improve near stall performance in a multistage compressor rig at NASA Glenn. Using specially designed stator vanes configured with internal actuation to deliver pulsating air through slots along the suction surface, a research study was performed to identify performance benefits using this microflow control approach. Pressure profiles and unsteady pressure measurements along the blade surface and at the shroud provided a dynamic look at the compressor during microflow air injection. These pressure measurements lead to a tracking algorithm to identify the onset of separation. The testing included steady air injection at various slot locations along the vane. The research also examined the benefit of pulsed injection and actively controlled air injection along the stator vane. Two types of actuation schemes were studied, including an embedded actuator for on-blade control. Successful application of an online detection and flow control scheme will be discussed. Testing showed dramatic performance benefit for flow reattachment and subsequent improvement in diffusion through the use of pulsed controlled injection. The paper will discuss the experimental setup, the blade configurations, and preliminary CFD results which guided the slot location along the blade. The paper will also show the pressure profiles and unsteady pressure measurements used to track flow control enhancement, and will conclude with the tracking algorithm for adjusting the control.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-2003-212356 , E-13944 , NAS 1.15:212356 , GT2003-38863 , Turbo Expo 2003; Jun 16, 2003 - Jun 19, 2003; Atlanta, GA; United States
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  • 3
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    The Effect of Ultrapolish on a Transonic Axial Rotor (2005)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Back-to-back testing has been done using NASA fan rotor 67 in the Glenn Research Center W8 Axial Compressor Test Facility. The rotor was baseline tested with a normal industrial RMS surface finish of 0.5-0.6 m (20-24 microinches) at 60, 80 and 100% of design speed. At design speed the tip relative Mach number was 1.38. The blades were then removed from the facility and ultrapolished to a surface finish of 0.125 m (5 microinch) or less and retested. At 100% speed near the design point, the ultrapolished blades showed approximately 0.3 - 0.5% increase in adiabatic efficiency. The difference was greater near maximum flow. Due to increased relative measurement error at 60 and 80% speed, the performance difference between the normal and ultrapolished blades was indeterminate at these speeds.
    Keywords: Mechanical Engineering
    Type: E-15193 , Turbo Expo 2005; Jun 06, 2005 - Jun 09, 2005; Reno, NV; United States
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  • 4
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    Experimental Results of the First Two Stages of an Advanced Transonic Core Compressor Under Isolated and Multi-Stage Conditions (2015)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: NASA's Environmentally Responsible Aviation (ERA) Program calls for investigation of the technology barriers associated with improved fuel efficiency of large gas turbine engines. Under ERA the task for a High Pressure Ratio Core Technology program calls for a higher overall pressure ratio of 60 to 70. This mean that the HPC would have to almost double in pressure ratio and keep its high level of efficiency. The challenge is how to match the corrected mass flow rate of the front two supersonic high reaction and high corrected tip speed stages with a total pressure ratio of 3.5. NASA and GE teamed to address this challenge by using the initial geometry of an advanced GE compressor design to meet the requirements of the first 2 stages of the very high pressure ratio core compressor. The rig was configured to run as a 2 stage machine, with Strut and IGV, Rotor 1 and Stator 1 run as independent tests which were then followed by adding the second stage. The goal is to fully understand the stage performances under isolated and multi-stage conditions and fully understand any differences and provide a detailed aerodynamic data set for CFD validation. Full use was made of steady and unsteady measurement methods to isolate fluid dynamics loss source mechanisms due to interaction and endwalls. The paper will present the description of the compressor test article, its predicted performance and operability, and the experimental results for both the single stage and two stage configurations. We focus the detailed measurements on 97 and 100 of design speed at 3 vane setting angles.
    Keywords: Aircraft Propulsion and Power
    Type: GRC-E-DAA-TN24306 , ASME Turbo Expo 2015; Jun 15, 2015 - Jun 19, 2015; Montreal, Quebec; Canada
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  • 5
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    Exprimental Results of the First Two Stages of an Advanced Transonic Core Compressor Under Isolated and Multi-Stage Conditions. (2015)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: NASA's Environmentally Responsible Aviation (ERA) Program calls for investigation of the technology barriers associated with improved fuel efficiency for large gas turbine engines. Under ERA, the highly loaded core compressor technology program attempts to realize the fuel burn reduction goal by increasing overall pressure ratio of the compressor to increase thermal efficiency of the engine. Study engines with overall pressure ratio of 60 to 70 are now being investigated. This means that the high pressure compressor would have to almost double in pressure ratio while keeping a high level of efficiency. NASA and GE teamed to address this challenge by testing the first two stages of an advanced GE compressor designed to meet the requirements of a very high pressure ratio core compressor. Previous test experience of a compressor which included these front two stages indicated a performance deficit relative to design intent. Therefore, the current rig was designed to run in 1-stage and 2-stage configurations in two separate tests to assess whether the bow shock of the second rotor interacting with the upstream stage contributed to the unpredicted performance deficit, or if the culprit was due to interaction of rotor 1 and stator 1. Thus, the goal was to fully understand the stage 1 performance under isolated and multi-stage conditions, and additionally to provide a detailed aerodynamic data set for CFD validation. Full use was made of steady and unsteady measurement methods to understand fluid dynamics loss source mechanisms due to rotor shock interaction and endwall losses. This paper will present the description of the compressor test article and its measured performance and operability, for both the single stage and two stage configurations. We focus the paper on measurements at 97% corrected speed with design intent vane setting angles.
    Keywords: Fluid Mechanics and Thermodynamics; Aircraft Propulsion and Power
    Type: GT2015-42727 , GRC-E-DAA-TN18595 , Turbo Technical Conference and Exposition; Jun 15, 2015 - Jun 19, 2015; Montreal, Quebec; Canada
    Format: application/pdf
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