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Hybrid electric aircraft propulsion is an emerging technology that presents a variety of potential benefits along with technical integration challenges. Developing these new propulsion architectures with their complex control systems, and ultimately proving their benefit, is a multistep process. This process includes concept development and analysis, dynamic simulation, hardware-in-the-loop testing, full-scale testing, and so on. This effort is being revolutionized and indeed enabled by new digital tools that support increasing the technology readiness level throughout the maturation process. As part of this Digital Transformation, NASA has developed a suite of publicly available digital tools that facilitate the path from concept to implementation. This paper describes the NASA-developed tools and puts them in the context of control system development for hybrid electric aircraft propulsion.

Aircraft icing remains a significant threat to aviation safety. Software that predicts the impingement and ice accretion on full aircraft geometries and aircraft components are in demand and NASA Glenn is committed to produce software that meets this need. One of the key parameters affecting an accurate prediction of iced geometry is the effect of ice roughness on the heat transfer coefficient. While many efforts have been made to implement the roughness in the flow solver, this report takes a correlation for roughness height distribution that is based on experimental measurements and demonstrates how to relate those measurements to an augmentation to the heat transfer coefficient provided by the flow solution. The outcome of this effort was the callibration of defaults for user supplied parameters to this correlation through comparison with 95 large glaze conditions from experiment by adjusting user-supplied parameters in the roughness augmentation equation.

Since, ice accretion can significantly degrade the performance and the stability of an airborne vehicle, it is imperative to be able to model it accurately. While ice accretion studies have been performed on airplane wings and helicopter blades in abundance, there are few that attempt to model the process on more complex geometries such as fuselages. This paper proposes a methodology that extends an existing in-house Extended Messinger solver to complex geometries by introducing the capability to work with unstructured grids and carry out spatial surface streamwise marching. For the work presented here commercial solvers such as STAR-CCM+ and ANSYS Fluent are used for the flow field and droplet dispersed phase computations. The ice accretion is carried out using an in-house icing solver called GT-ICE. The predictions by GT-ICE are compared to available experimental data, or to predictions by other solvers such as LEWICE and STAR-CCM+.

This paper describes an ice-crystal icing experiment conducted at the NASA Propulsion System Laboratory during June 2018. This test produced ice shape data on an airfoil for different test conditions similar to those inside the compressor region of a turbo-fan jet engine. Mixed-phase icing conditions were generated by partially freezing out a water spray using the relative humidity of flow as the primary parameter to control freeze-out. The paper presents the ice shape data and associated conditions which include pressure, velocity, temperature, humidity, total water content, melt ratio, and particle size distribution. The test featured a new instrument traversing system which allowed surveys of the flow and cloud. The purpose of this work was to provide experimental ice shape data and associated conditions to help develop and validate ice-crystal icing accretion models.

The formation of ice on aircraft is a highly dynamic process during which ice will expand and contract upon freezing and undergoing changes in temperature. Finite element analysis (FEA) simulations were performed investigating the stress/strain response of an idealized ice sample bonded to an acrylic substrate subjected to a uniform temperature change. The FEA predictions were used to guide the placement of strain gages on custom-built acrylic and aluminum specimens. Tee rosettes were placed in two configurations adjacent to thermocouple sensors. The specimens were then placed in icing conditions such that ice was grown on top of the specimen. It was hypothesized that the ice would expand on freezing and contract as the temperature of the interface returned to the equilibrium conditions.

An ice shape database has been created to document ice accretions on a 21-inch chord NACA0012 model and a 72-inch chord NACA 23012 airfoil model resulting from an exposure to a Supercooled Large Drop (SLD) icing cloud with a bimodal drop size distribution. The ice shapes created were documented with photographs, laser scanned surface measurements over a section of the model span, and measurement of the ice mass over the same section of each accretion. The icing conditions used in the test matrix were based upon previously used conditions on the same models but with an alternate approach to evaluation of drop distribution effects. Ice shapes resulting from the bimodal distribution as well as from equivalent monomodal drop size distributions were obtained and compared.

High Ice Water Content (HIWC) has been identified as a primary causal factor in numerous engine events over the past two decades. Previous attempts to develop a remote detection process utilizing modern commercial radars have failed to produce reliable results. This paper discusses the reasons for previous failures and describes a new technique that has shown very encouraging accuracy and range performance without the need for any modifications to industry’s current radar design(s). The performance of this new process was evaluated during the joint NASA/FAA HIWC RADAR II Flight Campaign in August of 2018. Results from that evaluation are discussed, along with the potential for commercial application, and development of minimum operational performance standards for future radar products.

Aerodynamic assessment of icing effects on swept wings is an important component of a larger effort to improve three-dimensional icing simulation capabilities. An understanding of ice-shape geometric fidelity and Reynolds and Mach number effects on iced-wing aerodynamics is needed to guide the development and validation of ice-accretion simulation tools. To this end, wind-tunnel testing was carried out for 8.9% and 13.3% scale semispan wing models based upon the Common Research Model airplane configuration. Various levels of geometric fidelity of an artificial ice shape representing a realistic glaze-ice accretion on a swept wing were investigated. The highest fidelity artificial ice shape reproduced all of the three-dimensional features associated with the glaze ice accretion. The lowest fidelity artificial ice shapes were simple, spanwise-varying horn ice geometries intended to represent the maximum ice thickness on the wing upper surface.

This paper describes a numerical model that simulates the thermal interaction between ice particles, water droplets, and the flowing air applicable during icing wind tunnel tests where there is significant phase-change of the cloud. It has been previously observed that test conditions, most notably temperature and humidity, change when the icing cloud is activated. It is hypothesized that the ice particles and water droplets thermally interact with the flowing air causing the air temperature and humidity to change by the time it reaches the test section. Unlike previous models where the air and particles are uncoupled, this model attempts to explain the observed changes in test conditions by coupling the conservation of mass and energy equations. The model is compared to measurements taken during wind tunnel tests simulating ice-crystal and mixed-phase icing that relate to ice accretions within turbofan engines.

In early 2015, a field campaign was conducted at the NASA Glenn Research Center in Cleveland, Ohio, USA. The purpose of the campaign is to test several prototype algorithms meant to detect the location and severity of in-flight icing (or icing aloft, as opposed to ground icing) within the terminal airspace. Terminal airspace for this project is currently defined as within 25 kilometers horizontal distance of the terminal, which in this instance is Hopkins International Airport in Cleveland. Two new and improved algorithms that utilize ground-based remote sensing instrumentation have been developed and were operated during the field campaign. The first is the ‘NASA Icing Remote Sensing System’, or NIRSS. The second algorithm is the ‘Radar Icing Algorithm’, or RadIA.

This paper will describe two recent modifications to the LEWICE software. The version described is under development and not ready for release. First, a capability for modeling ice crystals and mixed phase icing has been modified based on recent experimental data. Modifications have been made to the ice particle bouncing and erosion model. This capability has been added as part of a larger effort to model ice crystal ingestion in aircraft engines. Comparisons have been made to ice crystal ice accretions performed in the NRC Research Altitude Test Facility (RATFac). Second, modifications were made to the runback model based on data and observations from thermal scaling tests performed in the NRC Altitude Icing Tunnel. The runback model was modified to match film models used in the open literature. An empirical water shedding was also implemented. Comparisons were made to thermal deicing data taken at the NRC Altitude Icing Tunnel.

This paper presents measurements of ice accretion shape and surface temperature from ice-crystal icing experiments conducted jointly by the National Aeronautics and Space Administration (NASA) and the National Research Council (NRC) of Canada. The data comes from experiments performed at NRC's Research Altitude Test Facility (RATFac) in 2012. The measurements are intended to help develop models of the ice-crystal icing phenomenon associated with engine ice-crystal icing. Ice accretion tests were conducted using two different airfoil models (a NACA 0012 and wedge) at different velocities, temperatures, and pressures although only a limited set of permutations were tested. The wedge airfoil had several tests during which its surface was actively cooled. The ice accretion measurements included leading-edge thickness for both airfoils. The wedge and one case from the NACA 0012 model also included 2D cross-section profile shapes.

The formation of ice over lifting surfaces can affect aerodynamic performance. In the case of helicopters, this loss in lift and the increase in sectional drag forces will have a dramatic effect on vehicle performance. The ability to predict ice accumulation and the resulting degradation in rotor performance is essential to determine the limitations of rotorcraft in icing encounters. The consequences of underestimating performance degradation can be serious and so it is important to produce accurate predictions, particularly for severe icing conditions. The simulation of rotorcraft ice accretion is a challenging multidisciplinary problem that until recently has lagged in development over its counterparts in the fixed wing community. But now, several approaches for the robust coupling of a computational fluid dynamics code, a rotorcraft structural dynamics code and an ice accretion code have been demonstrated.

Three-dimensional simulations of the Honeywell ALF502 low pressure compressor (sometimes called a booster) using the NASA Glenn code GlennHT have been carried out. A total of eight operating points were investigated. These operating points are at, or near, points where engine icing has been determined to be likely. The results of this study were used, in a companion paper, for further analysis such as predicting collection efficiency of ice particles and ice growth rates at various locations in the compressor. In an effort to minimize computational effort, inviscid solutions with slip walls are produced. A mixing plane boundary condition is used between each blade row, resulting in convergence to steady state within each blade row. Comparisons of the results are made to other simplified analysis. An additional modification to the simulation process is also presented.

A flow and ice particle trajectory analysis was performed for the booster of the Honeywell ALF502 engine. The analysis focused on two closely related conditions one of which produced an icing event and another which did not during testing of the ALF502 engine in the Propulsion Systems Lab (PSL) at NASA Glenn Research Center. The flow analysis was generated using the NASA Glenn GlennHT flow solver and the particle analysis was generated using the NASA Glenn LEWICE3D v3.63 ice accretion software. The inflow conditions for the two conditions were similar with the main differences being that the condition that produced the icing event was 6.8 K colder than the non-icing event case and the inflow ice water content (IWC) for the non-icing event case was 50% less than for the icing event case.

The Propulsion Systems Laboratory (PSL) altitude test facility at NASA Glenn Research Center, has been used to test a full scale Honeywell turbofan engine at simulated altitude operating conditions. The PSL has spray bars to create a continuous cloud of fully glaciated ice crystals. The tests successfully duplicated the icing events that were experienced by the Honeywell engine (ALF502R-5) during flight through ice crystal clouds. After the ice cloud was turned on key engine performance parameters such as the fan speed, air flow rate, fuel flow rate, and compressor exit pressure and temperature responded immediately to the ingestion of the ice crystals. For some of the test points, these performance parameters remained unchanged from the initial response to the ice crystals, while during other test points the engine performance began to deteriorate to the point where an uncommanded loss of thrust control (engine rollback) was judged by the test engineers to have been imminent.

During the past two decades the occurrence of ice accretion within commercial high bypass aircraft turbine engines under certain operating conditions has been reported. Numerous engine anomalies have taken place at high altitudes that were attributed to ice crystal ingestion such as degraded engine performance, engine roll back, compressor surge and stall, and even flameout of the combustor. As ice crystals are ingested into the engine and low pressure compression system, the air temperature increases and a portion of the ice melts allowing the ice-water mixture to stick to the metal surfaces of the engine core. The focus of this paper is on estimating the effects of ice accretion on the low pressure compressor, and quantifying its effects on the engine system throughout a notional flight trajectory. In this paper it was necessary to initially assume a temperature range in which engine icing would occur.

Due to numerous engine power-loss events associated with high-altitude convective weather, ice accretion within an engine due to ice-crystal ingestion is being investigated. The National Aeronautics and Space Administration (NASA) and the National Research Council (NRC) of Canada are starting to examine the physical mechanisms of ice accretion on surfaces exposed to ice-crystal and mixed-phase conditions. In November 2010, two weeks of testing occurred at the NRC Research Altitude Facility utilizing a single wedge-type airfoil designed to facilitate fundamental studies while retaining critical features of a compressor stator blade or guide vane. The airfoil was placed in the NRC cascade wind tunnel for both aerodynamic and icing tests. Aerodynamic testing showed excellent agreement compared with CFD data on the icing pressure surface and allowed calculation of heat transfer coefficients at various airfoil locations.

In this work, a newly developed iced-aircraft modeling tool is applied to wings, engine inlets, and helicopter rotors. The tool is based on a multiscale-physics, unstructured finite-volume CFD approach and is applicable to general purpose aircraft icing applications. The present approach combines an Eulerian-based droplet-trajectory solver that is loosely coupled, in a time-accurate manner, to a surface-film and ice-evolution model. The goal of the model is to improve the fidelity of ice accretion modeling on dynamic geometries and for three-dimensional ice shapes typical of helicopter rotors. The numerical formulation is discussed and presented alongside 2D and 3D static validation cases, and dynamic helicopter rotors. The present results display good validation for predicting ice shape on a variety of geometries, and a strong initial capability of modeling ice forming on helicopters in forward flight.

To improve the understanding of the flow field within the NASA Glenn Icing Research Tunnel (IRT) with three different tunnel configurations, three-dimensional Reynold-Average Navier-Stokes (RANS) simulations were performed using the Menter-SST turbulence model. The 2000 tunnel configuration was simulated in the settling chamber from the spray bars to the test section. The 2009 tunnel configuration was simulated with vertical struts and multiple Mod-1 air jets implemented using embedded velocity profiles. The 2012 tunnel configuration has a new heat exchanger which was modeled starting from the exit of the heat exchanger to the test section. The results described herein focus on the flow turbulence since this defines test section performance but also is used to improve uniformity of the Liquid Water Content (LWC).