Results for “Compressor Modeling Modeling” — Page 30

NASA NTRS1980

Preliminary measurements of aerodynamic damping of a transonic compressor rotor

The aeroelastic behavior of a transonic compressor rotor operated in the MIT Blowdown Compressor Facility has been examined by means of piezoelectric motion sensors at the base of each of the 23 blades. Excitation has been observed due to rotating stall, due to an incipient flutter, and due to the facility startup transient. A method has been found for determining the aerodynamic damping force by modal analysis of the blade motion. Application of this technique to the example of excitation by rotating stall has led to the conclusions that the blade loading decreases in the stall cell, and that the damping force on the blades in the clean flow is in phase with blade velocity but opposite it in sign, leading to a logarithmic decrement of 0.2. This method of force derivation has quite general applicability as it requires only blade motion data such as are routinely acquired with strain gages. It is argued that models are needed for aerodynamic damping which focus on the effects of near neighbors of a given blade, since flutter often results in large response of isolated blades or small groups of blades.

Crawley, E. F.↗

NASA NTRS1989

A magnetic damper for first mode vibration reduction in multimass flexible rotors

Many rotating machines such as compressors, turbines and pumps have long thin shafts with resulting vibration problems, and would benefit from additional damping near the center of the shaft. Magnetic dampers have the potential to be employed in these machines because they can operate in the working fluid environment unlike conventional bearings. An experimental test rig is described which was set up with a long thin shaft and several masses to represent a flexible shaft machine. An active magnetic damper was placed in three locations: near the midspan, near one end disk, and close to the bearing. With typical control parameter settings, the midspan location reduced the first mode vibration 82 percent, the disk location reduced it 75 percent and the bearing location attained a 74 percent reduction. Magnetic damper stiffness and damping values used to obtain these reductions were only a few percent of the bearing stiffness and damping values. A theoretical model of both the rotor and the damper was developed and compared to the measured results. The agreement was good.

Kasarda, M. E. F.↗

NASA NTRS2001

Silicon Nitride Plates for Turbine Blade Application: FEA and NDE Assessment

Engine manufacturers are continually attempting to improve the performance and the overall efficiency of internal combustion engines. The thermal efficiency is typically improved by raising the operating temperature of essential engine components in the combustion area. This reduces the heat loss to a cooling system and allows a greater portion of the heat to be used for propulsion. Further improvements can be achieved by diverting part of the air from the compressor, which would have been used in the combustor for combustion purposes, into the turbine components. Such a process is called active cooling. Increasing the operating temperature, decreasing the cooling air, or both can improve the efficiency of the engine. Furthermore, lightweight, strong, tough hightemperature materials are required to complement efficiency improvement for nextgeneration gas turbine engines that can operate with minimum cooling. Because of their low-density, high-temperature strength, and thermal conductivity, ceramics are being investigated as potential materials for replacing ordinary metals that are currently used for engine hot section components. Ceramic structures can withstand higher operating temperatures and other harsh environmental factors. In addition, their low densities relative to metals helps condense component mass (ref. 1). The objectives of this program at the NASA Glenn Research Center are to develop manufacturing technology, a thermal barrier coating/environmental barrier coating (TBC/EBC), and an analytical modeling capability to predict thermomechanical stresses, and to do minimal burner rig tests of silicon nitride (Si3N4) and SiC/SiC turbine nozzle vanes under simulated engine conditions. Furthermore, and in support of the latter objectives, an optimization exercise using finite element analysis and nondestructive evaluation (NDE) was carried out to characterize and evaluate silicon nitride plates with cooling channels.

Abdul-Aziz, Ali↗

NASA NTRS1984

Preliminary flight test results of the F100 EMD engine in an F-15 airplane

A flight evaluation of the F100 Engine Model Derivative (EMD) is conducted. The F100 EMD is an advanced version of the F100 engine that powers the F15 and F16 airplanes. The F100 EMD features a bigger fan, higher temperature turbine, a Digital Electronic Engine Control system (DEEC), and a newly designed 16 segment afterburner, all of which results in a 15 to 20 percent increase in sea level thrust. The flight evaluations consist of investigation of performance (thrust, fuel flow, and airflow) and operability (transient response and airstart) in the F-15 airplane. The performance of the F100 EMD is excellent. Aircraft acceleration time to Mach 2.0 is reduced by 23 percent with two F100 EMD engines. Several anomalies are discovered in the operability evaluations. A software change to the DEEC improved the throttle, and subsequent Cooper Harper ratings of 3 to 4 are obtained. In the extreme upper left hand corner of the flight enveloped, compressor stalls occurr when the throttle is retarded to idle power. These stalls are not predicted by altitude facility tests or stability for the compressor.

Myers, L. P.↗

NASA NTRS2000

Modernization and Activation of the NASA Ames 11- by 11-Foot Transonic Wind Tunnel

The Unitary Plan Wind Tunnel (UPWT) was modernized to improve performance, capability, productivity, and reliability. Automation systems were installed in all three UPWT tunnel legs and the Auxiliaries facility. Major improvements were made to the four control rooms, model support systems, main drive motors, and main drive speed control. Pressure vessel repairs and refurbishment to the electrical distribution system were also completed. Significant changes were made to improve test section flow quality in the 11-by 11-Foot Transonic leg. After the completion of the construction phase of the project, acceptance and checkout testing was performed to demonstrate the capabilities of the modernized facility. A pneumatic test of the tunnel circuit was performed to verify the structural integrity of the pressure vessel before wind-on operations. Test section turbulence, flow angularity, and acoustic parameters were measured throughout the tunnel envelope to determine the effects of the tunnel flow quality improvements. The new control system processes were thoroughly checked during wind-off and wind-on operations. Manual subsystem modes and automated supervisory modes of tunnel operation were validated. The aerodynamic and structural performance of both the new composite compressor rotor blades and the old aluminum rotor blades was measured. The entire subsonic and supersonic envelope of the 11-by 11-Foot Transonic leg was defined up to the maximum total pressure.

Kmak, Frank J.↗

NASA NTRS2009

Large Eddy Simulation of Transonic Flow Field in NASA Rotor 37

The current paper reports on numerical investigations on the flow characteristics in a transonic axial compressor, NASA Rotor 37. The flow field was used previously as a CFD blind test case conducted by American Society of Mechanical Engineers in 1994. Since the CFD blind-test exercise, many numerical studies on the flow field in the NASA Rotor 37 have been reported. Although steady improvements have been reported in both numerical procedure and turbulence closure, it is believed that all the important aspects of the flow field have not been fully explained with numerical studies based on the Reynolds Averaged Navier-Stokes (RANS) solution. Experimental data show large dip in total pressure distribution near the hub at downstream of the rotor at 100% rotor speed. Most original numerical solutions from the blind test exercise did not predict this total pressure deficit correctly. This total pressure deficit at the rotor exit was attributed to a hub corner flow separation by the author. Several subsequent numerical studies with different turbulence closure model also calculated this dip in total pressure rise. Also, several studies attributed this total pressure deficit to a small leakage flow coming from the hub in the test article. As the experimental study cannot be repeated, either explanation cannot be validated. The primary purpose of the current investigation is to investigate the transonic flow field with both RANS and a Large Eddy Simulation (LES). The RANS approach gives similar results presented at the original blind test exercise. Although the RANS calculates higher overall total pressure rise, the total pressure deficit near the hub is calculated correctly. The numerical solution shows that the total pressure deficit is due to a hub corner flow separation. The calculated pressure rise from the LES agrees better with the measured total pressure rise especially near the casing area where the passage shock interacts with the tip clearance vortex and flow becomes unsteady due to this interaction. The LES simulation also calculates the total pressure rise deficit near the hub and it agrees well with the measured data.

Hah, Chunill↗

NASA NTRSDate not supplied

Transient Optimization of an Electrified Gas Turbine Engine Using Machine Learning

Gas turbine engines are designed with sufficient margin to prevent stall under normal operating conditions throughout their life. This compromise ensures that during rapid accelerations, compressor operation remains stable, but at the cost of efficiency and thrust responsiveness. The design margin encompasses multiple sources of uncertainty and systematic deviances from the operating line, the largest of which is the transient allowance. This set-aside accounts for the temporary incoordination of the engine spools during an acceleration while still enabling it to meet the certification requirement to accelerate from low to high power within a specified time, and without experiencing overtemperature, surge, stall, or other detrimental factors. Electrification of the powertrain provides the opportunity to address this reserve and truly optimize the design. The addition of electric machines inherent in hybrid propulsion concepts offers a means to interact with the engine shafts such that the necessary margin can be reduced, which can positively impact the engine design. By adjusting the amount of power extracted from or injected to the engine spools by the electric machines during transient operation, excursions from the operating line can be minimized. Past work using a dynamic engine model has shown that optimization of the fuel flow schedule during acceleration can reduce the required margin while still meeting the time requirement, and results are further improved when combined with power injection and extraction. The current work uses machine learning through a genetic algorithm to address the problem holistically by concurrently optimizing the electric machine power command and fuel flow acceleration schedule using an updated, higher fidelity version of the original engine model.

Stall Margin↗

NASA NTRSDate not supplied

Transient Optimization of an Electrified Gas Turbine Engine Using Machine Learning

Gas turbine engines are designed with sufficient margin to prevent stall under normal operating conditions throughout their life. This compromise ensures that during rapid accelerations, compressor operation remains stable, but at the cost of efficiency and thrust responsiveness. The design margin encompasses multiple sources of uncertainty and systematic deviances from the operating line, the largest of which is the transient allowance. This set-aside accounts for the temporary incoordination of the engine spools during an acceleration while still enabling it to meet the certification requirement to accelerate from low to high power within a specified time, and without experiencing overtemperature, surge, stall, or other detrimental factors. Electrification of the powertrain provides the opportunity to address this reserve and truly optimize the design. The addition of electric machines inherent in hybrid propulsion concepts offers a means to interact with the engine shafts such that the necessary margin can be reduced, which can positively impact the engine design. By adjusting the amount of power extracted from or injected to the engine spools by the electric machines during transient operation, excursions from the operating line can be minimized. Past work using a dynamic engine model has shown that optimization of the fuel flow schedule during acceleration can reduce the required margin while still meeting the time requirement, and results are further improved when combined with power injection and extraction. The current work uses machine learning through a genetic algorithm to address the problem holistically by concurrently optimizing the electric machine power command and fuel flow acceleration schedule using an updated, higher fidelity version of the original engine model.

Stall Margin↗

NASA NTRS2024

Transient Optimization of an Electrified Gas Turbine Engine Using Machine Learning

Gas turbine engines are designed with sufficient margin to prevent stall under normal operating conditions throughout their life. This compromise ensures that during rapid accelerations, compressor operation remains stable, but at the cost of efficiency and thrust responsiveness. The design margin encompasses multiple sources of uncertainty and systematic deviances from the operating line, the largest of which is the transient allowance. This set-aside accounts for the temporary incoordination of the engine spools during an acceleration while still enabling it to meet the certification requirement to accelerate from low to high power within a specified time, and without experiencing overtemperature, surge, stall, or other detrimental factors. Electrification of the powertrain provides the opportunity to address this reserve and truly optimize the design. The addition of electric machines inherent in hybrid propulsion concepts offers a means to interact with the engine shafts such that the necessary margin can be reduced, which can positively impact the engine design. By adjusting the amount of power extracted from or injected to the engine spools by the electric machines during transient operation, excursions from the operating line can be minimized. Past work using a dynamic engine model has shown that optimization of the fuel flow schedule during acceleration can reduce the required margin while still meeting the time requirement, and results are further improved when combined with power injection and extraction. The current work uses machine learning through a genetic algorithm to address the problem holistically by concurrently optimizing the electric machine power command and fuel flow acceleration schedule using an updated, higher fidelity version of the original engine model.

Stall Margin↗

NASA NTRS2006

Separation Control in a Multistage Compressor Using Impulsive Surface Injection

Control of flow separation using impulsive surface injection is investigated within the multistage environment of a low speed axial-flow compressor. Measured wake profiles behind a set of embedded stator vanes treated with suction-surface injection indicate significant reduction in flow separation at a variety of injection-pulse repetition rates and durations. The corresponding total pressure losses across the vanes reveal a bank of repetition rates at each pulse duration where the separation control remains nearly complete. This persistence allows for demands on the injected-mass delivery system to be economized while still achieving effective flow control. The response of the stator-vane boundary layers to infrequently applied short injection pulses is described in terms of the periodic excitation of turbulent strips whose growth and propagation characteristics dictate the lower bound on the band of optimal pulse repetition rates. The eventual falloff in separation control at higher repetition rates is linked to a competition between the benefits of pulse-induced mixing and the aggravation caused by the periodic introduction of low-momentum fluid. Use of these observations for impulsive actuator design is discussed and their impact on modeling the time-average effect of impulsive surface injection for multistage steady-flow simulation is considered.

Wundrow, David W.↗

NASA NTRS1998

Numerical Simulations of an Unsteady Rocket Launch from the AH-64D Apache Longbow Helicopter

Rocket and missile firings from attack helicopters can cause main engine compressor stall. Studies of this phenomenon suggest that the main engine ingests either the plume from the rockets or the rocket blast waves. This creates surges at the inlet face, causing a loss of power in the main engine. The objective of this project is to set-up a computational fluid dynamics (CFD) simulation of the AH-64D Apache Longbow helicopter during a rocket launch, in order to qualitatively study the fluid dynamics of the problem. This project presents a progression of three unsteady Navier-Stokes solutions. The first unsteady solution involves only a rocket launch from its launch canister. The second solution is a launch from a canister mounted on the Apache's wing-pylon assembly. The last solution includes the Apache main engine and fuselage. The computations use a series of structured, overset grid systems, which allow for a rocket moving in a prescribed path. The method implements a Roe upwind scheme with LU-SGS (lower-upper factored symmetric Gauss-Seidel). A rotor pressure disk model approximates the helicopter rotor, while the rocket engine exit properties are applied as a prescribed boundary condition. Although the project is only at the half-way point, the first and second CFD simulations suggest the possibility of pressure wave interference. Sudden surges in pressure occur from two sources: at rocket start-up, and as the rocket leaves the canister. Wave patterns set-up by these sources appear to propagate to the location of the engine inlet. However the simplified geometry simulation with the main engine needs to be performed before coming to a conclusion.

Okamoto, Kevin↗

NASA NTRS2013

Preliminary MIPCC Enhanced F-4 and F-15 Preformance Characteristics for a First Stage Reusable Launch Vehicle

Performance increases in turbojet engines can theoretically be achieved through Mass Injection Pre-Compressor Cooling (MIPCC), a process involving injecting water or oxidizer or both into an afterburning turbojet engine. The injection of water results in pre-compressor cooling, allowing the propulsion system to operate at high altitudes and Mach numbers. In this way, a MIPCC-enhanced turbojet engine could be used to power the first stage of a reusable launch vehicle or be integrated into an existing aircraft that could launch a 100-lbm payload to a reference 100-nm altitude orbit at 28 deg inclination. The two possible candidates for MIPCC flight demonstration that are evaluated in this study are the F-4 Phantom II airplane and the F-15 Eagle airplane (both of McDonnell Douglas, now The Boeing Company, Chicago, Illinois), powered by two General Electric Company (Fairfield, Connecticut) J79 engines and two Pratt & Whitney (East Hartford, Connecticut) F100-PW-100 engines, respectively. This paper presents a conceptual discussion of the theoretical performance of each of these aircraft using MIPCC propulsion techniques. Trajectory studies were completed with the Optimal Trajectories by Implicit Simulation (OTIS) software (NASA Glenn Research Center, Cleveland, Ohio) for a standard F-4 airplane and a standard F-15 airplane. Standard aircraft simulation models were constructed, and the thrust in each was altered in accordance with estimated MIPCC performance characteristics. The MIPCC and production aircraft model results were then reviewed to assess the feasibility of a MIPCC-enhanced propulsion system for use as a first-stage reusable launch vehicle; it was determined that the MIPCC-enhanced F-15 model showed a significant performance advantage over the MIPCC-enhanced F-4 model.

Air Breathing Launch Systems↗

NASA NTRS2013

Preliminary MIPCC Enhanced F-4 and F-15 Performance Characteristics for a First Stage Reusable Launch Vehicle

Performance increases in turbojet engines can theoretically be achieved through Mass Injection Pre-Compressor Cooling (MIPCC), a process involving injecting water or oxidizer or both into an afterburning turbojet engine. The injection of water results in pre-compressor cooling, allowing the propulsion system to operate at high altitudes and Mach numbers. In this way, a MIPCC-enhanced turbojet engine could be used to power the first stage of a reusable launch vehicle or be integrated into an existing aircraft that could launch a 100-lbm payload to a reference 100-nm altitude orbit at 28 deg inclination. The two possible candidates for MIPCC flight demonstration that are evaluated in this study are the F-4 Phantom II airplane and the F-15 Eagle airplane (both of McDonnell Douglas, now The Boeing Company, Chicago, Illinois), powered by two General Electric Company (Fairfield, Connecticut) J79 engines and two Pratt & Whitney (East Hartford, Connecticut) F100-PW-100 engines, respectively. This paper presents a conceptual discussion of the theoretical performance of each of these aircraft using MIPCC propulsion techniques. Trajectory studies were completed with the Optimal Trajectories by Implicit Simulation (OTIS) software (NASA Glenn Research Center, Cleveland, Ohio) for a standard F-4 airplane and a standard F-15 airplane. Standard aircraft simulation models were constructed, and the thrust in each was altered in accordance with estimated MIPCC performance characteristics. The MIPCC and production aircraft model results were then reviewed to assess the feasibility of a MIPCC-enhanced propulsion system for use as a first-stage reusable launch vehicle; it was determined that the MIPCC-enhanced F-15 model showed a significant performance advantage over the MIPCC-enhanced F-4 model.

MIPCC Reusable Launch Vehicle↗

NASA NTRS2010

A Unified Nonlinear Adaptive Approach for Detection and Isolation of Engine Faults

A challenging problem in aircraft engine health management (EHM) system development is to detect and isolate faults in system components (i.e., compressor, turbine), actuators, and sensors. Existing nonlinear EHM methods often deal with component faults, actuator faults, and sensor faults separately, which may potentially lead to incorrect diagnostic decisions and unnecessary maintenance. Therefore, it would be ideal to address sensor faults, actuator faults, and component faults under one unified framework. This paper presents a systematic and unified nonlinear adaptive framework for detecting and isolating sensor faults, actuator faults, and component faults for aircraft engines. The fault detection and isolation (FDI) architecture consists of a parallel bank of nonlinear adaptive estimators. Adaptive thresholds are appropriately designed such that, in the presence of a particular fault, all components of the residual generated by the adaptive estimator corresponding to the actual fault type remain below their thresholds. If the faults are sufficiently different, then at least one component of the residual generated by each remaining adaptive estimator should exceed its threshold. Therefore, based on the specific response of the residuals, sensor faults, actuator faults, and component faults can be isolated. The effectiveness of the approach was evaluated using the NASA C-MAPSS turbofan engine model, and simulation results are presented.

Tang, Liang↗

NASA NTRS2015

A Methodology to Assess the Capability of Engine Designs to Meet Closed-Loop Performance and Operability Requirements

Designing a closed-loop controller for an engine requires balancing trade-offs between performance and operability of the system. One such trade-off is the relationship between the 95 percent response time and minimum high-pressure compressor (HPC) surge margin (SM) attained during acceleration from idle to takeoff power. Assuming a controller has been designed to meet some specification on response time and minimum HPC SM for a mid-life (nominal) engine, there is no guarantee that these limits will not be violated as the engine ages, particularly as it reaches the end of its life. A characterization for the uncertainty in this closed-loop system due to aging is proposed that defines elliptical boundaries to estimate worst-case performance levels for a given control design point. The results of this characterization can be used to identify limiting design points that bound the possible controller designs yielding transient results that do not exceed specified limits in response time or minimum HPC SM. This characterization involves performing Monte Carlo simulation of the closed-loop system with controller constructed for a set of trial design points and developing curve fits to describe the size and orientation of each ellipse; a binary search procedure is then employed that uses these fits to identify the limiting design point. The method is demonstrated through application to a generic turbofan engine model in closed-loop with a simplified controller; it is found that the limit for which each controller was designed was exceeded by less than 4.76 percent. Extension of the characterization to another trade-off, that between the maximum high-pressure turbine (HPT) entrance temperature and minimum HPC SM, showed even better results: the maximum HPT temperature was estimated within 0.76 percent. Because of the accuracy in this estimation, this suggests another limit that may be taken into consideration during design and analysis. It also demonstrates the extension of the characterization to other attributes that contribute to the performance or operability of the engine. Metrics are proposed that, together, provide information on the shape of the trade-off between response time and minimum HPC SM, and how much each varies throughout the life cycle, at the limiting design points. These metrics also facilitate comparison of the expected transient behavior for multiple engine models.

systems analysis↗

NASA NTRS2015

A Methodology to Assess the Capability of Engine Designs to Meet Closed-loop Performance and Operability Requirements

Designing a closed-loop controller for an engine requires balancing trade-offs between performance and operability of the system. One such trade-off is the relationship between the 95% response time and minimum high-pressure compressor (HPC) surge margin (SM) attained during acceleration from idle to takeoff power. Assuming a controller has been designed to meet some specification on response time and minimum HPC SM for a mid-life (nominal) engine, there is no guarantee that these limits will not be violated as the engine ages, particularly as it reaches the end of its life. A characterization for the uncertainty in this closed-loop system due to aging is proposed that defines elliptical boundaries to estimate worst-case performance levels for a given control design point. The results of this characterization can be used to identify limiting design points that bound the possible con- troller designs yielding transient results that do not exceed specified limits in response time or minimum HPC SM. This characterization involves performing Monte Carlo simulation of the closed-loop system with controller constructed for a set of trial design points and developing curve fits to describe the size and orientation of each ellipse; a binary search procedure is then employed that uses these fits to identify the limiting design point. The method is demonstrated through application to a generic turbofan engine model in closed- loop with a simplified controller; it is found that the limit for which each controller was designed was exceeded by less than 4.76%. Extension of the characterization to another trade-off, that between the maximum high-pressure turbine (HPT) entrance temperature and minimum HPC SM, showed even better results: the maximum HPT temperature was estimated within 0.76%. Because of the accuracy in this estimation, this suggests another limit that may be taken into consideration during design and analysis. It also demonstrates the extension of the characterization to other attributes that contribute to the performance or operability of the engine. Metrics are proposed that, together, provide information on the shape of the trade-off between response time and minimum HPC SM, and how much each varies throughout the life cycle, at the limiting design points. These metrics also facilitate comparison of the expected transient behavior for multiple engine models.

engine control↗

NASA NTRS1996

Wave Engine Topping Cycle Assessment

The performance benefits derived by topping a gas turbine engine with a wave engine are assessed. The wave engine is a wave rotor that produces shaft power by exploiting gas dynamic energy exchange and flow turning. The wave engine is added to the baseline turboshaft engine while keeping high-pressure-turbine inlet conditions, compressor pressure ratio, engine mass flow rate, and cooling flow fractions fixed. Related work has focused on topping with pressure-exchangers (i.e., wave rotors that provide pressure gain with zero net shaft power output); however, more energy can be added to a wave-engine-topped cycle leading to greater engine specific-power-enhancement The energy addition occurs at a lower pressure in the wave-engine-topped cycle; thus the specific-fuel-consumption-enhancement effected by ideal wave engine topping is slightly lower than that effected by ideal pressure-exchanger topping. At a component level, however, flow turning affords the wave engine a degree-of-freedom relative to the pressure-exchanger that enables a more efficient match with the baseline engine. In some cases, therefore, the SFC-enhancement by wave engine topping is greater than that by pressure-exchanger topping. An ideal wave-rotor-characteristic is used to identify key wave engine design parameters and to contrast the wave engine and pressure-exchanger topping approaches. An aerodynamic design procedure is described in which wave engine design-point performance levels are computed using a one-dimensional wave rotor model. Wave engines using various wave cycles are considered including two-port cycles with on-rotor combustion (valved-combustors) and reverse-flow and through-flow four-port cycles with heat addition in conventional burners. A through-flow wave cycle design with symmetric blading is used to assess engine performance benefits. The wave-engine-topped turboshaft engine produces 16% more power than does a pressure-exchanger-topped engine under the specified topping constraints. Positive and negative aspects of wave engine topping in gas turbine engines are identified.

Welch, Gerard E.↗

NASA NTRS2017

A CFD Case Study of a Fan Stage with Split Flow Path Subject to Total Pressure Distortion Inflow

This report is the documentation of the work performed under the Hypersonic Project of the NASA's Fundamental Aeronautics Program. It was funded through Task Number NNC10E444T under GESS-2 Contract NNC06BA07B. The objective of the task is to develop advanced computational tools for the simulation of multi-stage turbomachinery in support of aeropropulsion. This includes work elements in extending the TURBO code and validating the multi-stage URANS (Unsteady Reynolds Averaged Navier Stokes) simulation results with the experimental data. The unsteady CFD (Computation Fluid Dynamics) calculations were performed in full wheel mode with and without screen generated total pressure distortion at the computational inflow boundary, as well as in single passage phase lag mode for uniform inflow. The experimental data were provided by NASA from the single stage RTA (Revolutionary Turbine Accelerator) fan test program.Significant non-uniform flow condition at the fan-face of the aeropropulsion system is frequentlyencountered in many of the advanced aerospace vehicles. These propulsion systems can be eithera podded or an embedded design employed in HWB (Hybrid Wing Body) airframe concept. It isalso a topic of interest in military applications, in which advanced air vehicles have already deployedsome form of embedded propulsion systems in their design because of the requirementsof compact and low observable inlets. Even in the conventional airframe/engine design, the fancould operate under such condition when the air vehicle is undergoing rapid maneuvering action.It is believed that a better understanding of the fan’s aerodynamic and aeromechanical responseto this type of operating condition or off design operation would be beneficial to designing distortiontolerant blades for improved engine operability.The objective for this research is to assess the capability of turbomachinery code as an analysistool in understanding the effects and evaluating the impact of flow distortion on the aerodynamicand aeromechanical performance of the fan in advanced propulsion systems. Results from thetesting of an advanced fan stage released by NASA are available and will be used here for CFDcode validation. The experiment was performed at NASA’s high speed compressor facility aspart of the RTA (Revolutionary Turbine Accelerator) demonstration project, a joint effort ofNASA Glenn Research Center and GE Aircraft Engines in developing an advanced Mach 4TBCC (Turbine Based Combined Cycle) turbofan/ramjet engine for access to space. Part of thetest was to assess the aerodynamic performance and operability of the fan stage under nonuniforminflow condition. Various flow distortion patterns were created at the fan-face by manipulatingsets of screens placed upstream of the wind tunnel. Measurements at the fan-face willprovide the necessary distortion flow information as the inflow boundary condition for the CFDin a full wheel simulation. Therefore the purpose of this work is to demonstrate the NASA supportedmulti-stage turbomachinery code, TURBO [1-5], in the aerodynamic performance analysisof a modern fan design operating under off design condition, and in particular to validate theCFD results with the RTA fan test data.A brief description of the RTA fan rig configuration is given in the next section, explaining onhow flow distortion were measured in the test and constructed for the CFD at the fan-face. It isfollowed by a section summarizing previous CFD work performed at NASA relevant to the currentfan configuration. A short description of the TURBO code is given next, followed by detailsin the computational model of the fan rig, the required computing resources, and the numericalprocedure for the simulations. The CFD results are presented in the discussion section and finallyconcluding remarks are summarized.

Fan Stage↗