TBCC Fan Stage Operability and Performance

NASA s Fundamental Aeronautics Program is investigating turbine-based 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. Studies performed under NASA s NGLT and the NASP High Speed Propulsion Assessment (HiSPA) program indicated a variable cycle turbofan/ramjet was the best configuration to satisfy access-to-space mission requirements because this configuration maximizes the engine thrust-to-weight ratio while minimizing frontal area. To this end, NASA and 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 (10X), fan speed (7X), inlet mass flow (3.5X), inlet pressure (8X), and inlet temperature (3X). The primary goal of the fan stage was to provide a high pressure ratio level with good efficiency at takeoff through the mid range of engine operation, while avoiding stall and losses at the higher flight Mach numbers, without the use of variable inlet guide vanes. Overall fan performance and operability therefore requires major consideration, as competing goals at different operating points and aeromechanical issues become major drivers in the design. To mitigate risk of meeting the unique design requirements for the fan stage, NASA and GE teamed to design and build a 57% engine scaled fan stage to be tested in NASA s transonic compressor facility. The objectives of this test are to assess the aerodynamic and aero mechanic performance and operability characteristics of the fan stage over the entire range of engine operation including: 1) sea level static take-off, 2) transition over large swings in fan bypass ratio, 3) transition from turbofan to ramjet, and 4) fan windmilling operation at high Mach flight conditions. In addition, the fan stage design was validated by performing pre-test CFD analysis using both GE proprietary and NASA s APNASA codes. Herein we will discuss 1) the fan stage design, 2) the experiment including the unique facility and instrumentation, and 3) the comparison of pre-test CFD analysis to initial aerodynamic test results for the baseline fan stage configuration. Measurements and pre-test analysis will be compared at 37%, 50%, 80%, 90%, and 100% of design speed to assess the ability of state-of-the-art design and analysis tools to meet the fan stage performance and operability requirements for turbine based propulsion for access to space

Experimental Investigation of the Flow Field in a Transonic, Axial Flow Compressor with Respect to the Development of Blockage and Loss

A detailed experimental investigation to understand and quantify the development of loss and blockage in the flow field of a transonic, axial flow compressor rotor has been undertaken. Detailed laser anemometer measurements were acquired upstream, within, and downstream of a transonic, axial compressor rotor operating at design and off-design conditions. The rotor was operated at 100%, 85%, 80%, and 60% of design speed which provided inlet relative Mach numbers at the blade tip of 1.48, 1.26, 1.18, and 0.89 respectively. At design speed the blockage is evaluated ahead of the rotor passage shock, downstream of the rotor passage shock, and near the trailing edge of the blade row. The blockage is evaluated in the core flow area as well as in the casing endwall region. Similarly at pm speed conditions for the cases of (1) where the rotor passage shock is much weaker than that at design speed and (2) where there is no rotor passage shock, the blockage and loss are evaluated and compared to the results at design speed. Specifically, the impact of the rotor passage shock on the blockage and loss development, pertaining to both the shock/boundary layer interactions and the shock/tip clearance flow interactions, is discussed. In addition, the blockage evaluated from the experimental data is compared to (1) an existing correlation of blockage development which was based on computational results, and (2) computational results on a limited basis. The results indicate that for this rotor the blockage in the endwall region is 2-3 times that of the core flow region and the blockage in the core flow region more than doubles when the shock strength is sufficient to separate the suction surface boundary layer. The distribution of losses in the care flow region indicate that the total loss is primarily comprised of the shock loss when the shock strength is not sufficient to separate the suction surface boundary layer. However, when the shock strength is sufficient to separate the suction surface boundary layer, the profile loss is comparable to the shock loss and can exceed the shock loss

NASA's Role in Gas Turbine Technology Development: Accelerating Technical Progress via Collaboration Between Academia, Industry, and Government Agencies

Given the maturity of the gas turbine engine since its invention and also considering the limited and flattened level of resources expected to be allocated for NASA aeronautics research and development, we ask the question are NASA technology investments still needed to enable future turbine engine-based propulsion systems? If so, what is NASAs unique role to justify NASAs investment? To address this topic, we will first review the accomplishments and the impact that NASA Glenn Research Center has made on turbine engine technologies over the last 78 years. Specifically, this paper discusses NASAs role and contributions to turbine engine development, specific to both 1) NASAs role in conducting experiments to understand flow physics and provide relevant benchmark validation experiments for Computational Fluid Dynamics (CFD) code development, validation, and assessment; and 2) the impact of technologies resulting from NASA collaborations with industry, academia, and other government agencies. Note that the scope of the discussion is limited to the NASA technology contributions with which the author was intimately associated, and does not represent the entirety of the NASA contributions to turbine engine technology. The specific research, development, and demonstrations discussed herein were selected to both 1) provide a comprehensive review and reference list of the technology and its impact, and 2) identify NASAs unique role and highlight how NASAs involvement resulted in additional benefit to the gas turbine engine community. Secondly, we will discuss current NASA collaborations that are in progress and provide a status of the results. Finally, we discuss the challenges anticipated for future turbine engine-based propulsion systems for civil aviation and identify potential opportunities for collaboration where NASA involvement would be beneficial. Ultimately, the gas turbine engine community will decide if NASA involvement is needed to contribute to the development of the design and analysis tools, databases, and technology demonstration programs to meet these challenges for future turbine engine-based propulsion systems

Effect of Variable Chord Length on Transonic Axial Rotor Performance Investigated

During the life of any gas turbine, blade erosion is present, especially for those units that are exposed to unfiltered air, such as aviation turbofan engines. The effect of this erosion is to reduce the blade chord progressively from the midspan to the tip region and to roughen and distort the blade surface. The effects of roughness on rotor performance have been documented by Suder et al. and Roberts. These papers indicate that the penalty for leading-edge roughness and erosion can be significant. Turbofan operators, therefore, restore chord length at routine maintenance intervals to regain performance before deterioration is too severe to salvage blades. As the rotor blades erode, the leading edge becomes rough - blunt and distorted from the nominal shape - and the aerodynamic performance suffers. Nominal performance can be recovered by recontouring the leading edges. This process, which inherently shortens the blade chord, can be used until the blade chord erodes to the stall limit. Below this chord length, which varies among engine-compressor types, a decrease of stall margin is likely. After compressor blade rework that includes leading edge recontouring, the blades have different chord lengths, ranging from blades that are near nominal chord length down to those near the stall chord limit. Furthermore, as blades erode below the stall limit, they must be replaced with new blades that have the full nominal chord length. Consequently, a set of compressor blades with varying chord lengths will be installed into each turbofan engine that goes through a complete maintenance cycle. The question arises, "Does fan or compressor performance depend on the order in which mixed-chord blades are installed into a fan or compressor disk?

Overview of the NASA Environmentally Responsible Aviation Project's Propulsion Technology Portfolio

The NASA Environmentally Responsible Aviation (ERA) Project is focused on developing and demonstrating integrated systems technologies to TRL 4-6 by 2020 that enable reduced fuel burn, emissions, and noise for futuristic air vehicles. The specific goals aim to simultaneously reduce fuel burn by 50%, reduce Landing and Take-off Nitrous Oxides emissions by 75% relative to the CAEP 6 guidelines, and reduce cumulative noise by 42 Decibels relative to the Stage 4 guidelines. These goals apply to the integrated vehicle and propulsion system and are based on a reference mission of 3000nm flight of a Boeing 777-200 with GE90 engines. This paper will focus primarily on the ERA propulsion technology portfolio, which consists of advanced combustion, propulsor, and core technologies to enable these integrated air vehicle systems goals. An overview of the ERA propulsion technologies will be described and the status and results to date will be presented

Experimental Study of Bypass Transition in a Boundary Layer

A detailed investigation to compare the boundary layer transition process in a low intensity disturbance environment to that in an environment in which the disturbances are initially non-linear in amplitude was conducted using a flat plate model. The transition mechanism based on linear growth of Tollmien Schlichting (T-S) waves was associated with a freestream turbulence level of 0.3 percent; however, for a freestream turbulence intensity of 0.65 percent and higher, the bypass transition mechanism prevailed. The results of detailed measurements acquired to study and compare the two transition mechanisms indicate that there exists a critical value for the peak rms of the velocity fluctuations within the boundary layer of approximately 3 to 3.5 percent of the freestream velocity. Once the unsteadiness within the boundary layer reached this critical value, turbulent bursting initiated, regardless of the transition mechanism. The two point correlations and simultaneous time traces within the transition region illustrate the features of a turbulent burst and its effect on the surrounding flowfield

Experimental and Computational Investigation of the Tip Clearance Flow in a Transonic Axial Compressor Rotor

Experimental and computational techniques are used to investigate tip clearance flows in a transonic axial compressor rotor at design and part speed conditions. Laser anemometer data acquired in the endwall region are presented for operating conditions near peak efficiency and near stall at 100% design speed and at near peak efficiency at 60% design speed. The role of the passage shock/leakage vortex interaction in generating endwall blockage is discussed. As a result of the shock/vortex interaction at design speed, the radial influence of the tip clearance flow extends to 20 times the physical tip clearance height. At part speed, in the absence of the shock, the radial extent is only 5 times the tip clearance height. Both measurements and analysis indicate that under part-speed operating conditions a second vortex, which does not originate from the tip leakage flow, forms in the endwall region within the blade passage and exits the passage near midpitch. Mixing of the leakage vortex with primary flow downstream of the rotor at both design and part speed conditions is also discussed

NASA Environmentally Responsible Aviation Projects Propulsion Technology Phase I Overview and Highlights of Accomplishments

The NASA Environmentally Responsible Aviation (ERA) Project is focused on developing and demonstrating integrated systems technologies to TRL 4-6 by 2020 that enable reduced fuel burn, emissions, and noise for futuristic air vehicles. The specific goals aim to simultaneously reduce fuel burn by 50%, reduce Landing and Take-off Nitrous Oxides emissions by 75% relative to the CAEP 6 guidelines, and reduce cumulative noise by 42 Decibels relative to the Stage 4 guidelines. These goals apply to the integrated vehicle and propulsion system and are based on a reference mission of 3000nm flight of a Boeing 777-200 with GE90 engines. This paper will focus primarily on the ERA propulsion technology portfolio, which consists of advanced combustion, propulsor, and core technologies to enable these integrated air vehicle systems goals. An overview of the ERA propulsion technologies will be described and highlights of the results obtained during the first phase of ERA will be presented

Laser anemometer measurements in a transonic axial-flow fan rotor

Laser anemometer surveys were made of the 3-D flow field in NASA rotor 67, a low aspect ratio transonic axial-flow fan rotor. The test rotor has a tip relative Mach number of 1.38. The flowfield was surveyed at design speed at near peak efficiency and near stall operating conditions. Data is presented in the form of relative Mach number and relative flow angle distributions on surfaces of revolution at nine spanwise locations evenly spaced from hub to tip. At each spanwise location, data was acquired upstream, within, and downstream of the rotor. Aerodynamic performance measurements and detailed rotor blade and annulus geometry are also presented so that the experimental results can be used as a test case for 3-D turbomachinery flow analysis codes

Design and performance of controlled-diffusion stator compared with original double-circular-arc stator

The capabilities of two stators, one with controlled-diffusion (CD) blade sections and one with double-circular-arc (DCA) blade sections, were compared. A CD stator was designed and tested that had the same chord length but half the blades of the DCA stator. The same fan rotor (tip speed, 429 m/sec; pressure ratio, 1.65) was used with each stator row. The design and analysis system is briefly described. The overall stage and rotor performances with each stator are compared, as are selected blade element data. The minimum overall efficiency decrement across the stator was approximately 1 percentage point greater with the CD balde sections than with the DCA blade sections