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NASA NTRS2017

Ice Crystal Icing Research at NASA

Ice crystals found at high altitude near convective clouds are known to cause jet engine power-loss events. These events occur due to ice crystals entering a propulsion system's core flowpath and accreting ice resulting in events such as uncommanded loss of thrust (rollback), engine stall, surge, and damage due to ice shedding. As part of a community with a growing need to understand the underlying physics of ice crystal icing, NASA has been performing experimental efforts aimed at providing datasets that can be used to generate models to predict the ice accretion inside current and future engine designs. Fundamental icing physics studies on particle impacts, accretion on a single airfoil, and ice accretions observed during a rollback event inside a full-scale engine in the Propulsion Systems Laboratory are summarized. Low fidelity code development using the results from the engine tests which identify key parameters for ice accretion risk and the development of high fidelity codes are described. These activities have been conducted internal to NASA and through collaboration efforts with industry, academia, and other government agencies. The details of the research activities and progress made to date in addressing ice crystal icing research challenges are discussed.

crysta↗

NASA NTRS2017

Ice Crystal Icing Research at NASA

Ice crystals found at high altitude near convective clouds are known to cause jet engine power-loss events. These events occur due to ice crystals entering a propulsion systems core flowpath and accreting ice resulting in events such as uncommanded loss of thrust (rollback), engine stall, surge, and damage due to ice shedding. As part of a community with a growing need to understand the underlying physics of ice crystal icing, NASA has been performing experimental efforts aimed at providing datasets that can be used to generate models to predict the ice accretion inside current and future engine designs. Fundamental icing physics studies on particle impacts, accretion on a single airfoil, and ice accretions observed during a rollback event inside a full-scale engine in the Propulsion Systems Laboratory are summarized. Low fidelity code development using the results from the engine tests which identify key parameters for ice accretion risk and the development of high fidelity codes are described. These activities have been conducted internal to NASA and through collaboration efforts with industry, academia, and other government agencies. The details of the research activities and progress made to date in addressing ice crystal icing research challenges are discussed.

icing↗

NASA NTRS2017

Ice Crystal Icing Research at NASA

Ice crystals found at high altitude near convective clouds are known to cause jet engine power-loss events. These events occur due to ice crystals entering a propulsion systems core flowpath and accreting ice resulting in events such as uncommanded loss of thrust (rollback), engine stall, surge, and damage due to ice shedding. As part of a community with a growing need to understand the underlying physics of ice crystal icing, NASA has been performing experimental efforts aimed at providing datasets that can be used to generate models to predict the ice accretion inside current and future engine designs. Fundamental icing physics studies on particle impacts, accretion on a single airfoil, and ice accretions observed during a rollback event inside a full-scale engine in the Propulsion Systems Laboratory are summarized. Low fidelity code development using the results from the engine tests which identify key parameters for ice accretion risk and the development of high fidelity codes are described. These activities have been conducted internal to NASA and through collaboration efforts with industry, academia, and other government agencies. The details of the research activities and progress made to date in addressing ice crystal icing research challenges are discussed.

upwind schemes↗

NASA NTRS2017

Engine Icing Data - An Analytics Approach

Engine icing researchers at the NASA Glenn Research Center use the Escort data acquisition system in the Propulsion Systems Laboratory (PSL) to generate and collect a tremendous amount of data every day. Currently these researchers spend countless hours processing and formatting their data, selecting important variables, and plotting relationships between variables, all by hand, generally analyzing data in a spreadsheet-style program (such as Microsoft Excel). Though spreadsheet-style analysis is familiar and intuitive to many, processing data in spreadsheets is often unreproducible and small mistakes are easily overlooked. Spreadsheet-style analysis is also time inefficient. The same formatting, processing, and plotting procedure has to be repeated for every dataset, which leads to researchers performing the same tedious data munging process over and over instead of making discoveries within their data. This paper documents a data analysis tool written in Python hosted in a Jupyter notebook that vastly simplifies the analysis process. From the file path of any folder containing time series datasets, this tool batch loads every dataset in the folder, processes the datasets in parallel, and ingests them into a widget where users can search for and interactively plot subsets of columns in a number of ways with a click of a button, easily and intuitively comparing their data and discovering interesting dynamics. Furthermore, comparing variables across data sets and integrating video data (while extremely difficult with spreadsheet-style programs) is quite simplified in this tool. This tool has also gathered interest outside the engine icing branch, and will be used by researchers across NASA Glenn Research Center. This project exemplifies the enormous benefit of automating data processing, analysis, and visualization, and will help researchers move from raw data to insight in a much smaller time frame.

Engine Icing↗

NASA NTRS2016

Preliminary Results From a Heavily Instrumented Engine Ice Crystal Icing Test in a Ground Based Altitude Test Facility

Preliminary results from the heavily instrumented ALF502R-5 engine test conducted in the NASA Glenn Research Center Propulsion Systems Laboratory are discussed. The effects of ice crystal icing on a full scale engine is examined and documented. This same model engine, serial number LF01, was used during the inaugural icing test in the Propulsion Systems Laboratory facility. The uncommanded reduction of thrust (rollback) events experienced by this engine in flight were simulated in the facility. Limited instrumentation was used to detect icing on the LF01 engine. Metal temperatures on the exit guide vanes and outer shroud and the load measurement were the only indicators of ice formation. The current study features a similar engine, serial number LF11, which is instrumented to characterize the cloud entering the engine, detect/characterize ice accretion, and visualize the ice accretion in the region of interest. Data were acquired at key LF01 test points and additional points that explored: icing threshold regions, low altitude, high altitude, spinner heat effects, and the influence of varying the facility and engine parameters. For each condition of interest, data were obtained from some selected variations of ice particle median volumetric diameter, total water content, fan speed, and ambient temperature. For several cases the NASA in-house engine icing risk assessment code was used to find conditions that would lead to a rollback event. This study further helped NASA develop necessary icing diagnostic instrumentation, expand the capabilities of the Propulsion Systems Laboratory, and generate a dataset that will be used to develop and validate in-house icing prediction and risk mitigation computational tools. The ice accretion on the outer shroud region was acquired by internal video cameras. The heavily instrumented engine showed good repeatability of icing responses when compared to the key LF01 test points and during day-to-day operation. Other noticeable observations are presented.

Enigine Icing↗

NASA NTRS2016

Preliminary Results From a Heavily Instrumented Engine Ice Crystal Icing Test in a Ground Based Altitude Test Facility

Preliminary results from the heavily instrumented ALF502R-5 engine test conducted in the NASA Glenn Research Center Propulsion Systems Laboratory are discussed. The effects of ice crystal icing on a full scale engine is examined and documented. This same model engine, serial number LF01, was used during the inaugural icing test in the Propulsion Systems Laboratory facility. The uncommanded reduction of thrust (rollback) events experienced by this engine in flight were simulated in the facility. Limited instrumentation was used to detect icing on the LF01 engine. Metal temperatures on the exit guide vanes and outer shroud and the load measurement were the only indicators of ice formation. The current study features a similar engine, serial number LF11, which is instrumented to characterize the cloud entering the engine, detect/ characterize ice accretion, and visualize the ice accretion in the region of interest. Data were acquired at key LF01 test points and additional points that explored: icing threshold regions, low altitude, high altitude, spinner heat effects, and the influence of varying the facility and engine parameters. For each condition of interest, data were obtained from some selected variations of ice particle median volumetric diameter, total water content, fan speed, and ambient temperature. For several cases the NASA in-house engine icing risk assessment code was used to find conditions that would lead to a rollback event. This study further helped NASA develop necessary icing diagnostic instrumentation, expand the capabilities of the Propulsion Systems Laboratory, and generate a dataset that will be used to develop and validate in-house icing prediction and risk mitigation computational tools. The ice accretion on the outer shroud region was acquired by internal video cameras. The heavily instrumented engine showed good repeatability of icing responses when compared to the key LF01 test points and during day-to-day operation. Other noticeable observations are presented.

turbomachinery↗

NASA NTRS2016

Preliminary Results From a Heavily Instrumented Engine Ice Crystal Icing Test in a Ground Based Altitude Test Facility

Preliminary results from the Heavily Instrumented ALF503R-5 Engine test conducted in the NASA Glenn Research Center Propulsion Systems Laboratory will be discussed. The effects of ice crystal icing on a full scale engine is examined and documented. This model engine, serial number LF01, was used during the inaugural icing test in the PSL facility. The reduction of thrust (rollback) events experienced by this engine in flight were replicated in the facility. Limited instrumentation was used to detect icing. Metal temperature on the exit guide vanes and outer shroud and the load measurement were the only indicators of ice formation. The current study features a similar engine, serial number LF11, which is instrumented to characterize the cloud entering the engine, detect characterize ice accretion, and visualize the ice accretion in the region of interest.

enigine icing↗

NASA NTRS2015

Complementary Aerodynamic Performance Datasets for Variable Speed Power Turbine Blade Section from Two Independent Transonic Turbine Cascades

Two independent experimental studies were conducted in linear cascades on a scaled, two-dimensional mid-span section of a representative Variable Speed Power Turbine (VSPT) blade. The purpose of these studies was to assess the aerodynamic performance of the VSPT blade over large Reynolds number and incidence angle ranges. The influence of inlet turbulence intensity was also investigated. The tests were carried out in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility and at the University of North Dakota (UND) High Speed Compressible Flow Wind Tunnel Facility. A large database was developed by acquiring total pressure and exit angle surveys and blade loading data for ten incidence angles ranging from +15.8deg to −51.0deg. Data were acquired over six flow conditions with exit isentropic Reynolds number ranging from 0.05×106 to 2.12×106 and at exit Mach numbers of 0.72 (design) and 0.35. Flow conditions were examined within the respective facility constraints. The survey data were integrated to determine average exit total-pressure and flow angle. UND also acquired blade surface heat transfer data at two flow conditions across the entire incidence angle range aimed at quantifying transitional flow behavior on the blade. Comparisons of the aerodynamic datasets were made for three "match point" conditions. The blade loading data at the match point conditions show good agreement between the facilities. This report shows comparisons of other data and highlights the unique contributions of the two facilities. The datasets are being used to advance understanding of the aerodynamic challenges associated with maintaining efficient power turbine operation over a wide shaft-speed range.

Cascade↗

NASA NTRS2015

Complementary Aerodynamic Performance Datasets for Variable Speed Power Turbine Blade Section from Two Independent Transonic Turbine Cascades

Two independent experimental studies were conducted in linear cascades on a scaled, two-dimensional mid-span section of a representative Variable Speed Power Turbine (VSPT) blade. The purpose of these studies was to assess the aerodynamic performance of the VSPT blade over large Reynolds number and incidence angle ranges. The influence of inlet turbulence intensity was also investigated. The tests were carried out in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility and at the University of North Dakota (UND) High Speed Compressible Flow Wind Tunnel Facility. A large database was developed by acquiring total pressure and exit angle surveys and blade loading data for ten incidence angles ranging from +15.8deg to −51.0deg. Data were acquired over six flow conditions with exit isentropic Reynolds number ranging from 0.05×106 to 2.12×106 and at exit Mach numbers of 0.72 (design) and 0.35. Flow conditions were examined within the respective facility constraints. The survey data were integrated to determine average exit total-pressure and flow angle. UND also acquired blade surface heat transfer data at two flow conditions across the entire incidence angle range aimed at quantifying transitional flow behavior on the blade. Comparisons of the aerodynamic datasets were made for three "match point" conditions. The blade loading data at the match point conditions show good agreement between the facilities. This report shows comparisons of other data and highlights the unique contributions of the two facilities. The datasets are being used to advance understanding of the aerodynamic challenges associated with maintaining efficient power turbine operation over a wide shaft-speed range.

Turbine↗

NASA NTRS2015

Comparison of Computational and Experimental Results for a Transonic Variable-speed Power-Turbine Blade Operating with Low Inlet Turbulence Levels

A computational assessment of the aerodynamic performance of the midspan section of a variable-speed power-turbine blade is described. The computation comprises a periodic single blade that represents the 2-D Midspan section VSPT blade that was tested in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. A commercial, off-the-shelf (COTS) software package, Pointwise and CFD++, was used for the grid generation and RANS and URANS computations. The CFD code, which offers flexibility in terms of turbulence and transition modeling options, was assessed in terms of blade loading, loss, and turning against test data from the transonic tunnel. Simulations were assessed at positive and negative incidence angles that represent the turbine cruise and take-off design conditions. The results indicate that the secondary flow induced at the positive incidence cruise condition results in a highly loaded case and transitional flow on the blade is observed. The negative incidence take-off condition is unloaded and the flow is very two-dimensional. The computational results demonstrate the predictive capability of the gridding technique and COTS software for a linear transonic turbine blade cascade with large incidence angle variation.

rotary wing↗

NASA NTRS2015

Simulation of VSPT Experimental Cascade Under High and Low Free-Stream Turbulence Conditions

Variable-Speed Power Turbines (VSPT) for rotorcraft applications operate at low Reynolds number and over a wide range in incidence associated with shaft speed change. A comprehensive linear cascade data set obtained includes the effects of Reynolds number, free-stream turbulence and incidence is available and this paper concerns itself with the presentation and numerical simulation of conditions resulting in a selected set of those data. As such, post-dictions of blade pressure loading, total-pressure loss and exit flow angles under conditions of high and low turbulence intensity for a single Reynolds number are presented. Analyses are performed with the three-equation turbulence models of Walters- Leylek and Walters and Cokljat. Transition, loading, total-pressure loss and exit angle variations are presented and comparisons are made with experimental data as available. It is concluded that at the low freestream turbulence conditions the Walters-Cokljat model is better suited to predictions while for high freestream conditions the two models generate similar predications that are generally satisfactory.

turbulence modeling↗

NASA NTRS2014

Aerodynamic Effects of High Turbulence Intensity on a Variable-Speed Power-Turbine Blade With Large Incidence and Reynolds Number Variations

The effects of high inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. These results are compared to previous measurements made in a low turbulence environment. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The current study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Assessing the effects of turbulence at these large incidence and Reynolds number variations complements the existing database. Downstream total pressure and exit angle data were acquired for 10 incidence angles ranging from +15.8deg to −51.0deg. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.12×10(exp 5) to 2.12×10(exp 6) and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 8 to 15 percent for the current study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitch/yaw probe located in a survey plane 7 percent axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the turbulence grid imposes.

Flegel, Ashlie B.↗

NASA NTRS2014

Aerodynamic Measurements of a Variable-Speed Power-Turbine Blade Section in a Transonic Turbine Cascade

The purpose of this thesis is to document the impact of incidence angle and Reynolds number variations on the three-dimensional flow field and midspan loss and turning of a two-dimensional section of a variable-speed power-turbine (VSPT) rotor blade. Aerodynamic measurements were obtained in a transonic linear cascade at NASA Glenn Research Center in Cleveland, Ohio. Steady-state data were obtained for 10 incidence angles ranging from +15.8deg to -51.0deg. At each angle, data were acquired at five flow conditions with the exit Reynolds number (based on axial chord) varying over an order-of-magnitude from 2.12×105 to 2.12×106. Data were obtained at the design exit Mach number of 0.72 and at a reduced exit Mach number of 0.35 as required to achieve the lowest Reynolds number. Midspan tota lpressure and exit flow angle data were acquired using a five-hole pitch/yaw probe surveyed on a plane located 7.0 percent axial-chord downstream of the blade trailing edge plane. The survey spanned three blade passages. Additionally, three-dimensional half-span flow fields were examined with additional probe survey data acquired at 26 span locations for two key incidence angles of +5.8deg and -36.7deg. Survey data near the endwall were acquired with a three-hole boundary-layer probe. The data were integrated to determine average exit total-pressure and flow angle as functions of incidence and flow conditions. The data set also includes blade static pressures measured on four spanwise planes and endwall static pressures.

Flegel, Ashlie B.↗

NASA NTRS2014

Simulation of VSPT Experimental Cascade Under High and Low Free-Stream Turbulence Conditions

Variable-Speed Power Turbines (VSPT) for rotorcraft applications operate at low Reynolds number and over a wide range in incidence associated with shaft speed change. A comprehensive linear cascade data set obtained includes the effects of Reynolds number, free-stream turbulence and incidence is available and this paper concerns itself with the presentation and numerical simulation of conditions resulting in a selected set of those data. As such, post-dictions of blade pressure loading, total-pressure loss and exit flow angles under conditions of high and low turbulence intensity for a single Reynolds number are presented. Analyses are performed with the three-equation turbulence models of Walters-Leylek and Walters and Cokljat. Transition, loading, total-pressure loss and exit angle variations are presented and comparisons are made with experimental data as available. It is concluded that at the low freestream turbulence conditions the Walters-Cokljat model is better suited to predictions while for high freestream conditions the two models generate similar predications that are generally satisfactory.

low pressure trubine↗

NASA NTRS2014

Aerodynamic Effects of High Turbulence Intensity on a Variable-Speed Power-Turbine Blade with Large Incidence and Reynolds Number Variations

The effects of high inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. These results are compared to previous measurements made in a low turbulence environment. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The current study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Assessing the effects of turbulence at these large incidence and Reynolds number variations complements the existing database. Downstream total pressure and exit angle data were acquired for 10 incidence angles ranging from +15.8deg to −51.0deg. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.12×10(exp 5) to 2.12×10(exp 6) and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 8 to 15 percent for the current study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitch/yaw probe located in a survey plane 7 percent axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the turbulence grid imposes.

Flegel, Ashlie B.↗