Joseph P. Veres

Mixed Phase Modeling in GlennICE with Application to Engine Icing

A capability for modeling ice crystals and mixed phase icing has been added to GlennICE. Modifications have been made to the particle trajectory algorithm and energy balance to model this behavior. This capability has been added as part of a larger effort to model ice crystal ingestion in aircraft engines. Comparisons have been made to four mixed phase ice accretions performed in the Cox icing tunnel in order to calibrate an ice erosion model. A sample ice ingestion case was performed using the Energy Efficient Engine (E 3 ) model in order to illustrate current capabilities. Engine performance characteristics were supplied using the Numerical Propulsion System Simulation (NPSS) model for this test case.

Axial and Centrifugal Compressor Mean Line Flow Analysis Method

This paper describes a method to estimate key aerodynamic parameters of single and multi -stage axial and centrifugal com pressors. Th is mean -line compressor code COMDES provides the capability of sizing single and multi -stage compressors quickly during the conceptual design process. Based on the compressible fluid flow equations and the Euler eq uation , the code can estimate rotor inlet and exit blade angles when run in the design mode. The design point rotor efficiency and stator losses are inputs to the code , and are modeled at off design. When run in the off -design analysis mode , it can be used to generate performance maps based on simple models for losses due to rotor incidence and inlet guide vane reset angle . The code can provide an improved understanding of basic aerodynamic parameters such as diffusion factor , loading levels and incidence, when matching multi -stage comp ressor blade rows at design and at part -speed operation. Rotor loading levels and relative velocity ratio are correlated to the onset of compressor surge. NASA Stage 37 and the three -stage NASA 74 -A axial compressors were analyzed and the results compared to test data. The code has been used to generate the performance map for the NASA 76 -B three -stage axial compressor featuring variable geometry. The compressor stages were aerodynamically matched at off -design speeds by adjusting the variable inlet guide v ane and variable stator geometry angles to control the rotor diffusion factor and incidence angles .

Modeling Commercial Turbofan Engine Icing Risk with Ice Crystal Ingestion

The occurrence of ice accretion within commercial high bypass aircraft turbine engines has been reported under certain atmospheric conditions. Engine anomalies have taken place at high altitudes that have been attributed to ice crystal ingestion, partially melting, and ice accretion on the compression system components. The result was degraded engine performance, and one or more of the following: loss of thrust control (roll back), compressor surge or stall, and flameout of the combustor. As ice crystals are ingested into the fan and low pressure compression system, the increase in air temperature causes a portion of the ice crystals to melt. It is hypothesized that this allows the ice-water mixture to cover the metal surfaces of the compressor stationary components which leads to ice accretion through evaporative cooling. Ice accretion causes a blockage which subsequently results in the deterioration in performance of the compressor and engine. The focus of this research is to apply an engine icing computational tool to simulate the flow through a turbofan engine and assess the risk of ice accretion. The tool is comprised of an engine system thermodynamic cycle code, a compressor flow analysis code, and an ice particle melt code that has the capability of determining the rate of sublimation, melting, and evaporation through the compressor flow path, without modeling the actual ice accretion. A commercial turbofan engine which has previously experienced icing events during operation in a high altitude ice crystal environment has been tested in the Propulsion Systems Laboratory (PSL) altitude test facility at NASA Glenn Research Center. The PSL has the capability to produce a continuous ice cloud which are ingested by the engine during operation over a range of altitude conditions. The PSL test results confirmed that there was ice accretion in the engine due to ice crystal ingestion, at the same simulated altitude operating conditions as experienced previously in flight. The computational tool was utilized to help guide a portion of the PSL testing, and was used to predict ice accretion could also occur at significantly lower altitudes. The predictions were qualitatively verified by subsequent testing of the engine in the PSL. The PSL test has helped to calibrate the engine icing computational tool to assess the risk of ice accretion. The results from the computer simulation identified prevalent trends in wet bulb temperature, ice particle melt ratio, and engine inlet temperature as a function of altitude for predicting engine icing risk due to ice crystal ingestion.

A Model to Assess the Risk of Ice Accretion due to Ice Crystal Ingestion in a Turbofan Engine and its Effects on Performance

4The occurrence of ice accretion within commercial high bypass aircraft turbine engines has been reported under certain atmospheric conditions. Engine anomalies have taken place at high altitudes that were attributed to ice crystal ingestion, partially melting, and ice accretion on the compression system components. The result was degraded engine performance, engine roll back, compressor surge and stall, and flameout of the combustor. As ice crystals are ingested into the fan and low pressure compression system, the air temperature increases and a portion of the ice crystals melt. This allows the ice-water mixture to stick to the metal surfaces of the compressor components. The resulting accretion causes a blockage on stationary components such as the stator vanes, and subsequently results in the deterioration in performance of the compressor and engine. The main focus of this research is the development of a computational tool that can estimate whether there is a risk of ice accretion by tracking key parameters through the compression system blade rows at all engine operating points within the flight trajectory. The tool has an engine system thermodynamic cycle code, coupled with a compressor flow analysis code, and an ice particle melt code that has the capability of determining the rate of sublimation, melting, and evaporation through the compressor blade rows. Assumptions are made to predict the complex physics involved in engine icing. Specifically, the code does not directly estimate ice accretion and does not have models for particle breakup, or erosion. Two key parameters have been suggested as conditions that must be met at the same location for ice accretion to occur: the local wet-bulb temperature to be near freezing and below, and the minimum local melt ratio must be above 10%. These parameters were deduced from analyzing normalized laboratory icing test data. These two parameters are the criteria that are used to determine whether ice accretion due to ice crystals is possible in an engine, and are used to identify the specific blade row where it could occur. Once the possibility of accretion is determined from these parameters, the degree of blockage due to ice accretion on the local stator vane can be estimated from an empirical model of ice growth rate and time spent at that operating point in the flight trajectory. The computational tool can be used to assess specific turbine engines to their susceptibility to ice accretion in an ice crystal environment.

Three Dimensional Simulation of Flow in an Axial Low Pressure Compressor at Engine Icing Operating Points

Viscous three-dimensional simulations of the Honeywell ALF502R-5 low pressure compressor (sometimes called a booster) using the NASA Glenn code GlennHT have been carried out. A total of ten simulations were produced. Five operating points are investigated, with each point run with two different wall thermal conditions. These operating points are at, or near, points where engine icing has been determined to be likely. In the future, the results of this study will be used for further analysis such as predicting collection efficiency of ice particles and ice growth rates at various locations in the compressor. A mixing plane boundary condition is used between each blade row, resulting in convergence to steady state within each blade row. The k-omega turbulence model of Wilcox, combined with viscous grid spacing near the wall on the order of one, is used to resolve the turbulent boundary layers. For each of the operating points, heat transfer coefficients are generated on the blades and walls. The heat transfer coefficients are produced by running the operating point with two different wall thermal conditions and then solving simultaneously for the heat transfer coefficient and adiabatic wall temperature at each point. Average Nusselt numbers are calculated for the most relevant surfaces. The values are seen to scale with Reynolds number to approximately a power of 0.7. Additionally, images of surface distribution of Nusselt number are presented. Qualitative comparison between the five operating points show that there is relatively little change in the character of the distribution. The dominant observed effect is that of an overall scaling, which is expected due to Reynolds number differences. One interesting aspect about the Nusselt number distribution is observed on the casing (outer diameter) downstream of the exit guide vanes (EGVs). The Nusselt number is relatively high between the pairs of EGVs, with two lower troughs downstream of each EGV trailing edge. This is of particular interest since rather complex ice shapes have been observed in that region. 1 Senior Research Engineer, Vantage Partners, LLC. 2 Research Scientist, Department of Mechanical and Aerospace Engineering, The Ohio State University. 3 Aerospace Engineer, Turbomachinery and Heat Transfer Branch, NASA Glenn Research Center. 4 Aerospace Engineer, Turbomachinery and Heat Transfer Branch, NASA Glenn Research Center. https://ntrs.nasa.gov/search.jsp?R=20170009137 2019-11-11T12:03:45+00:00Z

Engine Icing Modeling and Simulation (Part 2): Performance Simulation of Engine Rollback Phenomena

Abstract Ice buildup in the compressor section of a commercial aircraft gas turbine engine can cause a number of engine failures. One of these failure modes is known as engine rollback: an uncommanded decrease in thrust accompanied by a decrease in fan speed and an increase in turbine temperature. This paper describes the development of a model which simulates the system level impact of engine icing using the Commercial Modular Aero-Propulsion System Simulation 40k (C-MAPSS40k). When an ice blockage is added to C-MAPSS40k, the control system responds in a manner similar to that of an actual engine, and, in cases with severe blockage, an engine rollback is observed. Using this capability to simulate engine rollback, a proof-of-concept detection scheme is developed and tested using only typical engine sensors. This paper concludes that the engine control system’s limit protection is the proximate cause of iced engine rollback and that the controller can detect the buildup of ice particles in the compressor section. This work serves as a feasibility study for continued research into the detection and mitigation of engine rollback using the propulsion control system.

HIGH FIDELITY 3D TURBOFAN ENGINE SIMULATION WITH EMPHASIS ON TURBOMACHINERY-COMBUSTOR COUPLING

Introduction The 3-dimensional flow in the primary flow path of the GE90-94B high bypass ratio turbofan engine has been achieved. The simulation of the compressor components, the cooled high pressure turbine and the low pressure turbine was performed using the APNASA turbomachinery flow code. The combustor flow and chemistry were simulated using the National Combustor Code, NCC. The engine simulation matches the engine thermodynamic cycle for a sea-level takeoff condition. The fan, booster and OGV are corrected to the cycle condition from component simulations, whereas the high pressure compressor and turbines have been simulated at the cycle condition and coupled to the NCC code by passing profiles. Details of this coupling are presented. Significant gains in parallel computing are demonstrated which allow simulations to take place that can impact design. One of the goals of the Numerical Propulsion System Simulation (NPSS) Program at NASA Glenn Research Center has been to demonstrate a high-fidelity 3D Turbofan Engine Simulation. This simulation will support the multi-dimensional, multi-fidelity, multidiscipline concept of the design and analysis of propulsion systems for the future. This paper describes the current status of one major part of that goal: the complete turbofan engine simulation using an advanced 3-D Navier-Stokes turbomachinery solver, APNASA, coupled with the National Combustion Code, NCC. A production engine has been chosen for this demonstration: the GE90 turbofan engine shown in Fig. 1. A sea level, Mach 0.25, takeoff condition has been chosen for the simulation. The main reason is that detailed cooling flows for the turbine are well known at takeoff since this represents the cooled turbine design condition. Since the cooling flow represents a significant amount of the boundary condition information required for the simulation, it was felt this was a good point for the simulation. It also represents a condition where there are the highest temperatures and most stress in the engine, and is therefore a practical point to gain further understanding. _ C A S b A The GE90 development program included component testing of all the turbomachinery as well as the combustor. The full engine simulation effort has taken advantage of this. All the turbomachinery components 38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit 7-10 July 2002, Indianapolis, Indiana AIAA 2002-3769 Copyright

Modeling of Commercial Turbofan Engine With Ice Crystal Ingestion: Follow-On

The occurrence of ice accretion within commercial high bypass aircraft turbine engines has been reported under certain atmospheric conditions. Engine anomalies have taken place at high altitudes that have been attributed to ice crystal ingestion, partially melting, and ice accretion on the compression system components. The result was degraded engine performance, and one or more of the following: loss of thrust control (roll back), compressor surge or stall, and flameout of the combustor. As ice crystals are ingested into the fan and low pressure compression system, the increase in air temperature causes a portion of the ice crystals to melt. It is hypothesized that this allows the ice-water mixture to cover the metal surfaces of the compressor stationary components which leads to ice accretion through evaporative cooling. Ice accretion causes a blockage which subsequently results in the deterioration in performance of the compressor and engine. The focus of this research is to apply an engine icing computational tool to simulate the flow through a turbofan engine and assess the risk of ice accretion. The tool is comprised of an engine system thermodynamic cycle code, a compressor flow analysis code, and an ice particle melt code that has the capability of determining the rate of sublimation, melting, and evaporation through the compressor flow path, without modeling the actual ice accretion. A commercial turbofan engine which has previously experienced icing events during operation in a high altitude ice crystal environment has been tested in the Propulsion Systems Laboratory (PSL) altitude test facility at NASA Glenn Research Center. The PSL has the capability to produce a continuous ice cloud which is ingested by the engine during operation over a range of altitude conditions. The PSL test results confirmed that there was ice accretion in the engine due to ice crystal ingestion, at the same simulated altitude operating conditions as experienced previously in flight. The computational tool was utilized to help guide a portion of the PSL testing, and was used to predict ice accretion could also occur at significantly lower altitudes. The predictions were qualitatively verified by subsequent testing of the engine in the PSL. In a previous study, analysis of select PSL test data points helped to calibrate the engine icing computational tool to assess the risk of ice accretion. This current study is a continuation of that data analysis effort. The study focused on tracking the variations in wet bulb temperature and ice particle melt ratio through the engine core flow path. The results from this study have identified trends, while also identifying gaps in understanding as to how the local wet bulb temperature and melt ratio affects the risk of ice accretion and subsequent engine behavior.

Unsteady, Cooled Turbine Simulation Using a PC-Linux Analysis System

Summary The first stage of the high-pressure turbine (HPT) of the GE90 engine was simulated with a three-dimensional unsteady Navier-Sokes solver, MSU Turbo, which uses source terms to simulate the cooling flows. In addition to the solver, its pre-processor, GUMBO, and a post-processing and visualization tool, Turbomachinery Visual3 (TV3) were run in a Linux environment to carry out the simulation and analysis. The solver was run both with and without cooling. The introduction of cooling flow on the blade surfaces, case, and hub and its effects on both rotor-vane interac-tion as well as the effects on the blades themselves were the principle motivations for this study. The stud-ies of the cooling flow show the large amount of un-steadiness in the turbine and the corresponding hot streak migration phenomenon. This research on the GE90 turbomachinery has also led to a procedure for running unsteady, cooled turbine analysis on commod-ity PC’s running the Linux operating system. Introduction

Modeling of Highly Instrumented Honeywell Turbofan Engine Tested with Ice Crystal Ingestion in the NASA Propulsion System Laboratory

The Propulsion Systems Laboratory (PSL), an altitude test facility at NASA Glenn Research Center, has been used to test a highly instrumented turbine engine at simulated altitude operating conditions. This is a continuation of the PSL testing that successfully duplicated the icing events that were experienced in a previous engine (serial LF01) during flight through ice crystal clouds, which was the first turbofan engine tested in PSL. This second model of the ALF502R-5A serial number LF11 is a highly instrumented version of the previous engine. The PSL facility provides a continuous cloud of ice crystals with controlled characteristics of size and concentration, which are ingested by the engine during operation at simulated altitudes. Several of the previous operating points tested in the LF01 engine were duplicated to confirm repeatability in LF11. The instrumentation included video cameras to visually illustrate the accretion of ice in the low pressure compressor (LPC) exit guide vane region in order to confirm the ice accretion, which was suspected during the testing of the LF01. Traditional instrumentation included static pressure taps in the low pressure compressor inner and outer flow path walls, as well as total pressure and temperature rakes in the low pressure compressor region. The test data was utilized to determine the losses and blockages due to accretion in the exit guide vane region of the LPC. Multiple data points were analyzed with the Honeywell Customer Deck. A full engine roll back point was modeled with the Numerical Propulsion System Simulation (NPSS) code. The mean line compressor flow analysis code with ice crystal modeling was utilized to estimate the parameters that indicate the risk of accretion, as well as to estimate the degree of blockage and losses caused by accretion during a full engine roll back point. The analysis provided additional validation of the icing risk parameters within the LPC, as well as the creation of models for estimating the rates of blockage growth and losses.