Propulsion Systems

Single-Stage-to-Orbit Flow Path Studied
Successful Combustor Sector Tests Performed
Improved Ice-Scaling Method Demonstrated
Ice Measured on Multielement Wings
LEWICE Ice-Accretion Code Modified
Advanced Analytical Methods Developed for Spiral Bevel Gears
New Method Detects Gear Tooth Damage
High-Speed Exhaust Nozzle Test Proves Noise Goal Realistic
ASTOVL Lift Fan Nozzle Evaluated in Powered-Lift Facility
UTRC Tests Baseline SSTF Core Inlet/Bypass Nozzle
Small Two-Stage Axial Compressor Designed and Tested
Blade Row Interaction Effects on Flutter and Forced Response Modeled
First Phase of Wave Rotor Testing Completed
Test Bed for Active Control of Fan Noise Installed
Acoustic Data Acquisition Capability Improved in Low-Speed Wind Tunnel
Scale-Model, High-Bypass-Ratio Turbofans Tested at Simulated Takeoff/Approach

Because of the wide speed range over which they must operate and the multitude of takeoff and low-speed cycles available, propulsion system concepts for a single-stage-to-orbit (SSTO) vehicle have taken many forms in past studies. However, a single-flow-path arrangement would minimize both variable geometry and the number of auxiliary systems used over a limited speed range but necessarily carried to orbit. In this system a common propulsive duct accommodates the subsonic combustion "ramjet" cycle from about Mach 2 to Mach 6 and the supersonic combustion "scramjet" cycle from Mach 6 to its practical limit. (A rocket-based system is required for propulsion outside this envelope.) Design of the single duct satisfying these requirements must be biased toward efficient scramjet operation because the vehicle's gross takeoff weight is most sensitive to propulsion performance at higher speed. The design must, however, also be compatible with the ramjet and low-speed cycles that accelerate the vehicle through the transonic and supersonic flight regimes.

Single-stage-to-orbit flow-path model in wind tunnel.

An SSTO flow-path test program recently completed in the 1- by 1-Foot Supersonic Wind Tunnel investigated the compatibility of a flow path biased toward optimum scramjet performance with the ramjet mode of operation. The baseline geometry chosen for the study was that of the National AeroSpace Plane (NASP) concept demonstrator engine (CDE). The CDE configuration embodies the latest design philosophies of the NASP contractor team, and its use maximizes the benefit of the test results to the NASP program. The model was a 30%-scale, nonburning replica of the CDE. Although the model was not fueled, the entire ramjet compression process was properly modeled. The test was conducted at simulated flight Mach numbers of 3.5, 4.8, 6.3, and 7.1. Following completion of a baseline performance and operability matrix, the effects of other parameters, including forebody boundary layer height and isolator length, were determined. Performance with and without scramjet fuel injectors and inlet flap actuation struts was compared. The effect of inlet bleed was also studied.

The flow-path model was integrated into the sidewall of the tunnel test section. The inlet started and operated in a stable manner without bleed over the range of Mach numbers tested. The maximum stable contraction ratios exceeded those predicted for the CDE. Even though performance was less than that usually obtained with cruise, point-design configurations, these results are encouraging for the SSTO application.

Lewis contact: Charles J. Trefny, (216) 433-2162
Headquarters program office: OA

Successful Combustor Sector Tests Performed

NASA Lewis applies laser diagnostics and advanced probe diagnostic techniques to analyze details of combustion chemistry and uses flow visualization to optimize combustor fuel injector design and aerodynamics. Planar laser-induced fluorescence was successfully applied for the first time to a combustion flame tube at temperatures and pressures characteristic of actual combustor operating conditions. The images obtained suggest the degree of uniformity of the fuel injection process, allowing fuel injector designers to optimize their injectors for more uniform fuel distribution in the combustor.

In other experiments fuel droplet sizes and velocities were obtained in the combustion flame tube at actual engine operating pressures using water to simulate fuel. This difficult measurement in a extremely dense spray was obtained using a phase Doppler anemometer, which simultaneously measures droplet sizes and velocities nonintrusively. The combustion and spray diagnostics were obtained through unique high-pressure and high-temperature windows that provide optical access on all four sides of the square flame-tube test section.

In another breakthrough in diagnostic techniques a gas chromatograph/mass spectrometry system was applied to gas samples extracted through ultraclean probes inserted into the combustion test section. The breakthrough was in obtaining parts-per-billion resolution of hydrocarbon concentrations using industry-standard equipment but performing special dilution techniques to increase the resolution capability.

Lewis contact: Richard W. Niedzwiecki, (216) 433-3407
Headquarters program office: OA

Improved Ice-Scaling Method Demonstrated

Determination of improved ice-scaling method.

For tests in an icing tunnel, scaling methods must produce ice shapes with the same features and relative size for the scaled test as for the full-size test. A number of scaling "laws" have been derived over the past 40 years or so, but all have limitations. As part of an ongoing in-house program, recent work at NASA Lewis has demonstrated an improved scaling method for testing when the cloud water content is different from that desired. The new method reproduces both shape and quantity of ice accurately. This result permits greater use of relatively low-cost tunnel tests in place of flight tests and permits more reliable extrapolation of flight data to different cloud conditions.

Lewis contact: Dr. David N. Anderson, (216) 433-3585
Headquarters program office: OA

Ice on multielement wing in Lewis Icing Research Tunnel.

In a cooperative venture between NASA Lewis and McDonnell Douglas, ice shapes have been measured on a multielement airfoil in the Lewis Icing Research Tunnel. This venture is part of an ongoing NASA Lewis program to study advanced-design airfoils in icing conditions. Ice-shape measurements on a two-dimensional airfoil were completed in July 1994. These shapes are being machined in aluminum and applied to wing sections for performance testing in the NASA Langley Low Turbulence Pressure Tunnel in September 1994. These performance tests will be the first of their kind with ice shapes. NASA Langley and McDonnell Douglas have been testing multielement airfoils without ice for several years. The results from the clean-airfoil tests will be compared with the September test results using simulated ice to determine the performance penalty due to ice.

Lewis contact: Dr. Jaiwon Shin, (216) 433-8714
Headquarters program office: OA

LEWICE Ice-Accretion Code Modified

LEWICE had been found in the past to predict ice shapes very accurately for many operating conditions. However, for some cases involving long icing times, the ice shapes were not considered to be reliable. As part of an on-going effort the numerical methods used in LEWICE have been improved during the past year to enhance the accuracy of the ice-shape predictions. A new version of LEWICE that overcomes many of the previous problems has been made available to the aeronautics industry. We are working to improve the accuracy of LEWICE by developing better physical models of the processes involved in ice accretion.

The three-dimensional version of the ice-accretion code, called LEWICE3D, was originally structured to calculate ice accretion on swept wings, such as those on high-speed aircraft. This code has been modified to permit calculations for engine inlets as well. In addition, the ability to interact with various types of user-supplied codes has been incorporated to enhance the code's usefulness. LEWICE3D with these modifications is being used to develop data for the icing certification of the Gulfstream G4 aircraft.

Lewis contacts: William B. Wright, (216) 977-1246; Colin S. Bidwell, (216) 433-3947
Headquarters program office: OA

Advanced Analytical Methods Developed for Spiral Bevel Gears

Model used for three-dimensional contact of spiral bevel gears.

A thermal analysis procedure was developed that uses a finite element technique to conduct a full steady-state and transient three-dimensional heat transfer analysis of meshing spiral bevel gears. Industry now uses a method developed in the 1960's that gives a single value of the "flash temperature," or the maximum temperature attained during the meshing process. Designers typically compare this value with their successful applications where thermal problems were avoided or use it to compare design options. The new procedure can predict the full three-dimensional temperature field and the "flash temperature" behavior of the entire meshing gear.

Another new analytical technique was developed for full three-dimensional contact modeling during the meshing process. This nonlinear finite element analysis code has successfully modeled multiple teeth in contact. The advantage of this technique is that the load distribution and contact are found and not assumed for the multiple contacts that can occur in this gear system. This analytical technique can predict load sharing, transmission error, bending stress, and contact stresses.

Bibliography

Handschuh, R.; and Kicher, T.: A Method for Thermal Analysis of Spiral Bevel Gears. NASA TM-106612, ARL-TR-457, 1994.
Bibel, G.; Kumar, A.; and Tiku, K.: Contact Analysis of Spiral Bevel Gears. NASA CR-195305, ARL-CR-146, 1994.

Lewis contact: Robert F. Handschuh, (216) 433-3969
Headquarters program office: OA

New Method Detects Gear Tooth Damage

The new method, NA4*, was developed to detect the onset of gear tooth damage and to continue to react to the damage as it increases. It was designed to overcome the deficiencies of other gear fault detection methods, which either fail to react for certain damage modes or tend to lose their effectiveness as the damage increases. NA4*, a time-based, nondimensional statistical parameter, uses the vibration signal from an accelerometer mounted on the housing of a gear system. The vibration signal is first time averaged, using a time pulse synchronous to the rotation of the gear being monitored, before method NA4* is applied. The time-synchronous averaging eliminates all vibration not coincident with the rotation of the gear being evaluated. NA4* produces a value of 3 under normal conditions for all gear types regardless of the transmission system configuration. NA4* has the ability to compare the current condition of the gear being monitored to a weighted baseline of the gear in "good" condition, which allows it to both detect faults and grow with the increasing damage.

Face gear run 5 results.

Method NA4* was applied to experimental vibration data from a number of spur gear, face gear, and spiral bevel gear fatigue tests conducted on various fatigue test rigs at NASA Lewis and involving a variety of damage modes. The primary purpose of these test rigs is to study the effects of gear tooth design, gear materials, and lubrication types on the fatigue strength of aircraft-quality gears. Gear damage modes ranged from moderate pitting to complete fracture of several teeth on one gear. NA4* robustly detected a wide range of gear tooth damage modes and magnitudes on a variety of gear types. In a face gear fatigue test NA4* increased from the nominal value of 3 to a value of 150, which reflected the major damage of two fractured teeth on the gear. In a spur gear test, NA4* increased to a value of 9 as pitting started and finally to a value of 45 as pitting covered the complete face width of the damaged gear tooth.

Method NA4* can be a significant part of an onboard health and usage monitoring system for current and future rotorcraft. The early and reliable detection of gear tooth damage is a crucial part of an in-flight fault warning system.

Bibliography

Decker, H.J.; et al.: An Enhancement to the NA4 Gear Vibration Diagnostic Parameter. NASA TM-106553, 1994.
Zakrajsek, J.J.; et al.: An Analysis of Gear Fault Detection Methods as Applied to Pitting Fatigue Failure Data. NASA TM-105950, 1993.
Zakrajsek, J.J.; et al.: Application of Fault Detection Techniques to Spiral Bevel Gear Fatigue Data. NASA TM-106467, 1994.
Zakrajsek, J.J.: A Review of Transmission Diagnostics Research at NASA Lewis Research Center. NASA TM-106746, 1994.

Lewis contacts: James J. Zakrajsek, (216) 433-3968; Harry J. Decker, (216) 433-3964
Headquarters program office: OA

High-Speed Exhaust Nozzle Test Proves Noise Goal Realistic

HSR exhaust nozzle in CE-22 test facility.

NASA Lewis has recently completed a joint experimental program with Pratt & Whitney and General Electric to measure the internal performance of several candidate HSR exhaust nozzle concepts. The models were tested at simulated altitude conditions in the NASA Lewis CE-22 facility to determine their performance over a wide range of geometric variations and nozzle pressure ratios representing conditions from takeoff to supersonic cruise.

The tests were successfully completed and results agreed well with contractors' predictions, indicating that the high-speed nozzle performance goal was realistic. In addition, improvements to the CE-22 facility resulted in 99.75% accuracy and repeatability in both thrust and flow measurements, enabling the results of these tests to establish a broad comprehensive data base. The data base will be used to validate current performance prediction methodology and to help select an optimal exhaust nozzle design for the HSCT propulsion system.

Lewis contact: David W. Lam, (216) 433-8875
Headquarters program office: OA

ASTOVL Lift Fan Nozzle Evaluated in Powered-Lift Facility

Nozzle vectored at degrees to augment thrust durng takeoff.

Allison identified the Powered-Lift Facility (PLF) at NASA Lewis as an ideal facility for such a test, and Lewis subsequently entered into a cooperative test program with them. The PLF is a large static thrust stand with a six-component, multiaxis thrust-measuring system. It is located inside a 130-ft-diameter geodesic dome, which shields against noise during testing. Allison provided a one-third-scale model of the lift fan nozzle for testing in PLF. Nozzle flow rates, thrust forces, and vector angles were measured for various nozzle orientations and pressure ratios. The nozzle is designed to vector up to 10°ree; for yaw and to pitch forward up to 20°ree; or rearward up to 60°ree; to augment forward thrust during takeoff.

Results from the test demonstrated that the nozzle met design requirements and also agreed well with the thrust and flow performance predicted by Allison's CFD analyses. The test was performed in a timely manner, enabling Allison to proceed as scheduled with fabrication of the large-scale lift fan hardware for the NASA Ames test.

Lewis contact: David W. Lam, (216) 433-8875
Headquarters program office: OA

UTRC Tests Baseline SSTF Core Inlet/Bypass Nozzle

UTRC began work on the SSTF core inlet/bypass nozzle by screening several potential concepts. The concepts had to provide the weight flow required by the engine compressor, high inlet pressure recovery, and high nozzle thrust performance. They selected one concept, with a variable core inlet and a variable bypass nozzle, for further study with the GASP Navier-Stokes code. The GASP results indicated that high inlet pressure recovery was possible with significant bleed of the boundary layer entering the core inlet. UTRC then began testing a two-dimensional scale model of the selected concept in their new dual-flow test facility.

Core inlet/bypass nozzle installed in UTRC dual-flow test facilitiy.

The dual-flow facility was constructed to permit testing of the variable core inlet/variable bypass nozzle at flight conditions ranging from Mach 0.9 (loiter) to Mach 2.8 (design point). The core inlet was designed with variable geometry cowls to control the inlet captured airflow and contraction ratio. In addition, the inlet has a bleed scoop to remove the low-energy boundary layer along the hub. The bypass nozzle was designed with a variable-geometry flap to optimize nozzle performance at various operating conditions.

Initial tests of the core inlet/bypass nozzle model installed in the dual-flow test facility have been successfully completed. Fully supersonic flow in the core inlet and bypass nozzle was achieved and maintained as the cowls of the core inlet were translated. Additional tests will be conducted to characterize the performance of the core inlet and bypass nozzle and to determine any necessary design modifications. Finally, experimental results will be compared with the flow predictions obtained previously with the GASP code.

Lewis contact: Nicholas J. Georgiadis, (216) 433-3958
Headquarters program office: OA

Small Two-Stage Axial Compressor Designed and Tested

NASA Lewis, Allison, and the U.S. Army established a joint program to design a small, highly loaded, two-stage axial compressor using an advanced three-dimensional, viscous, multistage flow analysis code. We would try to demonstrate the performance improvements achievable through the use of such modeling (ref. 1). We would also use the test data to validate and calibrate the multistage code used in the design phase.

Two-stage rotor assembly for advanced small turboshaft compressor.

The performance testing has been completed. The compressor achieved a 5:1 pressure ratio in two stages, and the performance data indicated good efficiency and stall margin. The test provides a unique data base for small multistage axial compressors. Even though the stall margin was acceptable, we have established a follow-on program to enhance the original compressor design.

Reference

1. Adamczyk, J.J.: Model Equations for Simulating Flows in Multistage Turbomachinery. ASME Paper 85-GT-226, 1985.

Lewis contact: Dr. David P. Miller, (216) 433-8352
Headquarters program office: OA

Blade Row Interaction Effects on Flutter and Forced Response Modeled

Rotor satbility with and without aerodynamic coupling.

NASA Lewis has developed a procedure for calculating intra-blade-row, unsteady aerodynamic interactions that relies on results from isolated-blade-row unsteady aerodynamic analyses. Using influence coefficients that express the unsteady forces on one blade row due to the motion of another, we obtained an aeroelastic model that accounts for coupling of the vibratory responses of multiple blade rows. The model was applied to two configurations, each consisting of three blade rows. The flutter analysis showed that interaction effects can be destabilizing, and the forced-response analysis showed that interaction effects can significantly increase the resonant response of a blade row.

Bibliography

Buffum, D.H.: Blade Row Interaction Effects on Flutter and Forced Response. NASA TM-106438, 1993.

Lewis contact: Dr. Daniel H. Buffum, (216) 433-3759
Headquarters program office: OA

First Phase of Wave Rotor Testing Completed

Wave rotor sensitivity to leakage losses was determined by installing false end-walls, whose position relative to the rotor could be altered to vary the leakage gap. Passage gradual opening time and viscous losses were characterized by parametrically varying rotor length, passage width, and rotor speed. As the passages were made shorter and wider, the efficiency improved, indicating that friction plays a major role.

The design effort is under way for the second phase of the research, which involves a four-port rotor. The four-port cycle, along with a conventional external combustor, is appropriate for use as a topping cycle. The experiment, however, is designed to use a heater rather than a combustor, to keep temperatures lower.

Wave rotor performance.

Extensive validation of the one-dimensional (ref. 1) and two-dimensional (ref. 2) codes used for modeling the wave rotor dynamics has also been accomplished this year. The one-dimensional design/analysis code was used to design the four-port wave rotor cycle and to generate wave rotor component maps. The maps were used by the Advanced Aeronautics Office in a cycle study to evaluate wave rotor performance enhancement in a GE-90 class, an Allison 250, and a Lycoming T-55 class engine. The studies show that small turboshaft engines are significantly enhanced, both in increased specific power and in reduced specific fuel consumption, by a wave rotor topping cycle. Finally, in collaboration with NASA Lewis, the Allison Engine Co. is investigating the application of a wave rotor to an existing engine, using available hardware.

References

1. Paxson, D.E.: A Comparison Between Numerically Modelled and Experimentally Measured Loss Mechanisms in Wave Rotors. AIAA Paper 93-2522, 1993.
   
2. Welch, G.: Two-Dimensional Numerical Study of Wave Rotor Flow Dynamics. AIAA Paper 93-2525, 1993.

Test Bed for Active Control of Fan Noise Installed

A large low-speed fan has been designed, fabricated, and installed in the NASA Lewis Aeroacoustic Propulsion Laboratory for the purpose of developing and demonstrating various ANC concepts aimed at reducing fan tone noise. This work is part of the Advanced Subsonic Technology Noise Reduction Program. Several concepts identified by Pratt & Whitney, General Electric, Hersh Acoustical Engineering, and NASA will be tested on this 48-in.-diameter fan as a proof of concept before a full-scale engine demonstration. The first test of an ANC system is scheduled for the first half of 1995.

Active noise control fan installed in Aeroacoustic Propulsion Laboratory.

The combination of the large fan diameter and low (400 ft/sec) tip speed produces fan tones of the same frequencies produced by full-size advanced engines. A unique feature of this fan is the direct attachment of the rotor centerbody to the rig support column, eliminating the need for struts or pylons, which could contaminate acoustic measurements. Another important feature is the built-in rotating microphones, which separate the fan tone into individual modes, allowing better determination of the noise generation mechanisms.

Lewis contact: Laurence J. Heidelberg, (216) 433-3859
Headquarters program office: OA

Acoustic Data Acquisition Capability Improved in Low-Speed Wind Tunnel

These changes have reduced the measured in-flow background noise in the test section at Mach 0.20 by about 7 to 8 dB at frequencies from 200 Hz to 40 kHz. Far-field acoustic instrumentation in the tunnel consists of fixed, wall-mounted microphones and a remote-controlled translating microphone probe. The translating probe covers most of the axial length of the tunnel test section. The computer data acquisition system also moves the microphone to (typically) 48 data positions, pausing for data acquisition at each position.

9- by 15-Foot Low-Speed Wind Tunnel test section showing fixed and translating microphones.

Tunnel background noise calibration tests with the fixed, wall-mounted microphones have shown that upstream microphone holder wakes may interact with downstream microphones. In particular, the microphone holders generate a viscous wake with less than a 7°ree; half-angle and also an apparent tip vortex that persists directly downstream from the probe. Thus, the wall microphones must be displaced vertically and into the flow to minimize wake interaction. The translating microphone probe avoids the issue of upstream wake interaction while providing significantly more data measurement locations.

Comparisons of fixed and translating microphone data for a model turbofan show good agreement at comparable tunnel locations. These results suggest that fixed, wall-mounted microphones are not required for model fan noise tests. The translating microphone probe was shown to take accurate noise samples with good spatial resolution and eliminates possible wake disturbances from upstream microphones.

Lewis contact: Richard P. Woodward, (216) 433-3923
Headquarters program office: OA

Scale-Model, High-Bypass-Ratio Turbofans Tested at Simulated Takeoff/Approach

The GE UPS is designed to simulate full-scale engine components in scale-model size. The complete engine fan module (comprising engine nacelle, bypass fan stage, and nozzle), the first one-and-a-half stages of the engine core booster stage, the wing mounting pylon, and engine acoustic treatment within the nacelle are all simulated. The 22-in.-diameter test model fan is powered by a four-stage air turbine drive module using high-pressure, high-temperature (450 psi, 550 °ree;F) air at flow rates to 25 lb/sec. During the test approximately 1900 performance and operating parameters were obtained, both for measuring the fan module performance and for monitoring the safety and health of the entire model. Unique with this model is the ability to measure fan module performance by using force balances mounted internally in the model. Fan thrust and torque forces can be measured by using a two-component rotating balance; bypass stator and nacelle drag and torque forces can be measured by using a six-component static balance. Acoustic measurements of the model were made by using an array of 13 wall microphones at fixed azimuthal locations, in addition to a unique track-mounted, axially translating microphone with the capability to stop at any selected angular location along the entire length of the test section.

GE universal propulsion simulator installed in wind tunnel.

The data obtained during this test will be used by NASA and GE research engineers to validate the many computer design and analysis codes used to design both high-bypass-ratio fan stage components and fan module acoustic treatment. The data will be compared with the predictions generated by engineers using computer codes to determine how well scale-model engine components can be designed to simulate full-scale engine performance and where these codes can be improved for better engine component performance predictions in the future.

Lewis contacts: Christopher E. Hughes, (216) 433-3924;
Richard P. Woodward, (216) 433-3923
Headquarters program office: OA

Last updated 1995

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