Scott A. Thorp

Compressor Stability Enhancement Using Discrete Tip Injection

Mass injection upstream of the tip of a high-speed axial compressor rotor is a stability enhancement approach known to be effective in suppressing stall in tip-critical rotors. This process is examined in a transonic axial compressor rotor through experiments and time-average Navier-Stokes CFD simulations. Measurements and simulations for discrete injection are presented for a range of injection rates and distributions of injectors around the annulus.The simulations indicate that tip injection increases stability by unloading the rotor tip and that increasing injection velocity improves the effectiveness of tip injection. For the tested rotor, experimental results demonstrate that at 70% speed the stalling flow coefficient can be reduced by 30% using an injected massflow equivalent to 1% of the annulus flow. At design speed, the stalling flow coefficient was reduced by 6% using an injected massflow equivalent to 2% of the annulus flow. The experiments show that stability enhancement is related to the mass-averaged axial velocity at the tip. For a given injected massflow, the mass averaged axial velocity at the tip is increased by injecting flow over discrete portions of the circumference as opposed to full-annular injection. The implications of these results on the design of recirculating casing treatments and other methods to enhance stability will be discussed.Copyright

Compressor Stall Control Through Endwall Recirculation

Experiments that demonstrate the use of endwall recirculation to control the stall of transonic compressor stages are described. Endwall recirculation of a compressor stage is implemented by bleeding air from the casing downstream of a stator blade row and injecting the air as a wall jet upstream of a preceding rotor blade row. The bleed ports, injection ports, and recirculation channels are circumferentially discrete, and occupy only 20–30% of the circumference. The development of compact wall-jet injectors is described first. Next, the results of proof-of-concept steady recirculation tests on a single-stage transonic compressor are presented. Finally, the potential for using endwall recirculation to increase the stability of transonic highly-loaded multistage compressors is demonstrated through results from a rig test of simulated recirculation driving both a steady injected flow and an unsteady injected flow commanded by closed-loop active control during compressor operation at 78–100% of design speed. In this test air from an external source was injected upstream of several rotor blade rows while compressor bleed was increased by an amount equivalent to the injected massflow. During closed loop control, wall static pressure fluctuations were monitored and the injected flow rate was controlled to reduce the stalling mass flow. The use of wall jet injection to study the dynamics of transonic compressor stages is also discussed.Copyright

The Effect of Variable Chord Length on Transonic Axial Rotor Performance

Aircraft fan and compressor blade leading edges suffer from atmospheric particulate erosion that reduces aerodynamic performance. Recontouring the blade leading edge region can restore blade performance. This process typically results in blades of varying chord length. The question therefore arises as to whether performance of refurbished fans and compressors could be further improved if blades of varying chord length are installed into the disk in a certain order. To investigate this issue the aerodynamic performance of a transonic compressor rotor operating with blades of varying chord length was measured in back-to-back compressor test rig entries. One half of the rotor blades were the full nominal chord length while the remaining half of the blades were cut back at the leading edge to 95% of chord length and recontoured. The rotor aerodynamic performance was measured at 100, 80, and 60% of design speed for three blade installation configurations: nominal-chord blades in half of the disk and short-chord blades in half of the disk; four alternating quadrants of nominal-chord and short-chord blades; nominal-chord and short-chord blades alternating around the disk. No significant difference in performance was found between configurations, indicating that blade chord variation is not important to aerodynamic performance above the stall chord limit if leading edges have the same shape. The stall chord limit for most civil aviation turbofan engines is between 94-96% of nominal (new) blade chord.

Testing and Performance Verification of a High Bypass Ratio Turbofan Rotor in an Internal Flow Component Test Facility

A 1/5 scale model rotor representative of a current technology, high bypass ratio, turbofan engine was installed and tested in the W8 single-stage, high-speed, compressor test facility at NASA Glenn Research Center (GRC). The same fan rotor was tested previously in the GRC 9x15 Low Speed Wind Tunnel as a fan module consisting of the rotor and outlet guide vanes mounted in a flight-like nacelle. The W8 test verified that the aerodynamic performance and detailed flow field of the rotor as installed in W8 were representative of the wind tunnel fan module installation. Modifications to W8 were necessary to ensure that this internal flow facility would have a flow field at the test package that is representative of flow conditions in the wind tunnel installation. Inlet flow conditioning was designed and installed in W8 to lower the fan face turbulence intensity to less than 1.0 percent in order to better match the wind tunnel operating environment. Also, inlet bleed was added to thin the casing boundary layer to be more representative of a flight nacelle boundary layer. On the 100 percent speed operating line the fan pressure rise and mass flow rate agreed with the wind tunnel data to within 1 percent. Detailed hot film surveys of the inlet flow, inlet boundary layer and fan exit flow were compared to results from the wind tunnel. The effect of inlet casing boundary layer thickness on fan performance was quantified. Challenges and lessons learned from testing this high flow, low static pressure rise fan in an internal flow facility are discussed.

The Effect of Ultrapolish on a Transonic Axial Rotor

Back-to-back testing was 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 root-mean-square (RMS) surface finish of 0.5 μm to 0.6 μm (20 microinches to 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 μm (5 microinch) or less and retested. At 100% speed near the design point, the ultrapolished blades showed approximately 0.3% to 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.

Results of an Advanced Fan Stage Operating Over a Wide Range of Speed and Bypass Ratio: Part I—Fan Stage Design and Experimental Results

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, 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 multipoint operation to accommodate variations in bypass ratio (10X), fan speed (7X), inlet mass flow (3.5X), inlet pressure (8X), and inlet temperature (3X). 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% corrected design speed. Whereas, in the companion paper [1], we will provide an assessment of NASA’s APNASA code’s ability to predict the fan stage performance & operability over a wide range of speed and bypass ratio.Copyright