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A research program was conducted to evaluate the effectiveness of icing tunnel hybrid model design. A hybrid design is where the full-scale leading edge of a wing section is maintained only to a certain percentage of the local chord, while the aft section of the model is redesigned into a shortened or truncated planform. An initial study was conducted in 2020 where the ice shape geometries on a full-chord length version of the swept CRM65 wing model were compared to those from the hybrid version of CRM65 that were obtained in the NASA Icing Research Tunnel in 2015. The results were reported in a 2021 paper. For most test conditions, the overall size and shape of the ice shapes compared well. However, the ice shapes from the full-chord model were generally slightly smaller than those from the hybrid model.

The Glenn Icing Computational Environment (GlennICE) is a computational tool designed to calculate ice growth on complex three-dimensional geometries using the input from a user-supplied computational fluid dynamics (CFD) solution for the geometry of interest. The most significant developments in the advancement of GlennICE have been investigating the convergence of the collection efficiency, efficiently finding trajectories, and improving the refinement methodology. Such developments have increased the efficiency of GlennICE for practical engineering application. With the increasing demand for applying GlennICE for more memory-intensive problems, the scalability of GlennICE has yet to be investigated. This paper is aimed at presenting a method to benchmark the scalability of GlennICE utilizing a relevant engineering problem within a parallel environment.

The Icing Research Tunnel at NASA Glenn follows the recommended practice for calibration outlined in SAE’s ARP5905. The calibration team has followed the schedule of a full calibration every five years with a check calibration done every six months following. The liquid water content of the IRT has maintained stability within the stated specifications of variation within +/- 10% of the curve fit equation generated from calibration data. Using past measurements and data trends, IRT characterization engineers wanted to develop methods for the ability to know when data were not within variation. Trends can be observed in the liquid water content measurement process by constructing statistical process control charts. This paper describes data processing procedures for the Multi-Element Sensor in the IRT, including collision efficiency corrections, canonical correlation analysis, process for rejection of data, and construction of control charts.

When examining the literature on the adhesion strength of impact ice, there have been a wide range of methodologies tried to measure the required stresses to induce interfacial delamination. Utilizing the Icing Research Tunnel at the NASA Glenn Research Center to generate the impact ice required for this work, several different mechanical tests have been and are being developed to determine the stresses along the interface between ice and coupon. This set of tests includes the technical mature modified lap joint test which has been used to conduct ice adhesion studies through a wide sweep of icing conditions. To conduct in situ ice adhesion measurements inside of the Icing Research Tunnel, several new experiments are currently being developed to make ice adhesion measurements during and immediately after ice accretion.

Numerical solutions have been generated which simulate flow inside an aircraft engine flying at altitude through an ice crystal cloud. The geometry used for this study is the Honeywell Uncertified Research Engine (HURE) which was recently tested in the NASA Propulsion Systems Laboratory (PSL) in January 2018. The simulations were carried out at predicted operating points with a potential risk of ice accretion. The extent of the simulation is from upstream of the engine inlet to downstream past the strut in the core and bypass. The flow solution is produced using GlennHT, a NASA in-house code. A mixing plane approximation is used upstream and downstream of the fan. The use of the mixing plane allows for steady state solutions in the relative frame. The flow solution is then passed on to LEWICE3D for particle trajectory, impact and breakup prediction. The LEWICE3D code also uses a mixing plane approximation at the boundaries upstream and downstream of the fan.

This paper reports on temperature and humidity measurements from a series of ice-crystal icing tunnel experiments conducted in June 2018 at the Propulsion Systems Laboratory at the NASA Glenn Research Center. The tests were fundamental in nature and were aimed at investigating the icing processes on a two-dimensional NACA0012 airfoil subjected to artificially generated icing clouds. Prior to the tests on the airfoil, a suite of instruments, including total temperature and humidity probes, were used to characterize the thermodynamic flow and icing cloud conditions of the facility. Two different total temperature probes were used in these tests which included a custom designed rearward facing probe and a commercial self-heating total temperature probe. The rearward facing probe, the main total temperature probe, is being designed to reduce and mitigate the contaminating effects of icing and ingestion of ice crystals and water droplets at the probe’s inlet.

Computational ice shapes were generated on the boundary layer ingesting engine nacelle of the D8 Double Bubble aircraft. The computations were generated using LEWICE3D, a well-known CFD icing post processor. A 50-bin global drop diameter discretization was used to capture the collection efficiency due to the direct impingement of water onto the engine nacelle. These discrete results were superposed in a weighted fashion to generate six drop size distributions that span the Appendix C and O regimes. Due to the presence of upstream geometries, i.e. the fuselage nose, the trajectories of the water drops are highly complex. Since the ice shapes are significantly correlated with the collection efficiency, the upstream fuselage nose has a significant impact on the ice accretion on the engine nacelle. These complex trajectories are caused by the ballistic nature of the particles and are thus exacerbated as particle size increases.

NASA and the FAA conducted two flight campaigns to quantify onboard weather radar measurements with in-situ measurements of high concentrations of ice crystals found in deep convective storms. The ultimate goal of this research was to improve the understanding of high ice water content (HIWC) and develop onboard weather radar processing techniques to detect regions of HIWC ahead of an aircraft to enable tactical avoidance of the potentially hazardous conditions. Both HIWC RADAR campaigns utilized the NASA DC-8 Airborne Science Laboratory equipped with a Honeywell RDR-4000 weather radar and in-situ microphysical instruments to characterize the ice crystal clouds. The purpose of this paper is to summarize how these campaigns were conducted and highlight key results. The first campaign was conducted in August 2015 with a base of operations in Ft. Lauderdale, Florida.

This paper analyzes ice crystal icing accretion data and evaluates a thermodynamic ice crystal icing model, which has been previously presented, to describe the possible mechanisms of icing within the core of a turbofan jet engine. The model functions between two distinct ice accretions based on a surface energy balance: freeze-dominated icing and melt-dominated icing. Freeze-dominated icing occurs when liquid water (from melted ice crystals) freezes and accretes on a surface along with the existing ice of the impinging water and ice mass. This freeze-dominated icing is characterized as having strong adhesion to the surface. The amount of ice accretion is partially dictated by a freeze fraction, which is the fraction of impinging liquid water that freezes. Melt-dominated icing occurs as unmelted ice on a surface accumulates. This melt-dominated icing is characterized by weakly bonded surface adhesion.

The aerodynamic effects of Cold Soaked Fuel Frost have become increasingly significant as airworthiness authorities have been asked to allow it during aircraft take-off. The Federal Aviation Administration and the Finnish Transport Safety Agency signed a Research Agreement in aircraft icing research in 2015 and started a research co-operation in frost formation studies, computational fluid dynamics for ground de/anti-icing fluids, and de/anti-icing fluids aerodynamic characteristics. The main effort has been so far on the formation and aerodynamic effects of CSFF. To investigate the effects, a generic high-lift common research wind tunnel model and DLR-F15 airfoil, representing the wing of a modern jet aircraft, was built including a wing tank cooling system. Real frost was generated on the wing in a wind tunnel test section and the frost thickness was measured with an Elcometer gauge. Frost surface geometry was measured with laser scanning and photogrammetry.

Artificial ice shapes of various geometric fidelity were tested on a wing model based on the Common Research Model. Low Reynolds number tests were conducted at Wichita State University’s Walter H. Beech Memorial Wind Tunnel utilizing an 8.9% scale model, and high Reynolds number tests were conducted at ONERA’s F1 wind tunnel utilizing a 13.3% scale model. Several identical geometrically-scaled ice shapes were tested at both facilities, and the results were compared at overlapping Reynolds and Mach numbers. This was to ensure that the results and trends observed at low Reynolds number could be applied and continued to high, near-flight Reynolds number. The data from Wichita State University and ONERA F1 agreed well at matched Reynolds and Mach numbers. The lift and pitching moment curves agreed very well for most configurations.

The National Aeronautics and Space Administration (NASA) conducted a full scale ice crystal icing turbofan engine test using an obsolete Allied Signal ALF502-R5 engine in the Propulsion Systems Laboratory (PSL) at NASA Glenn Research Center. The test article used was the exact engine that experienced a loss of power event after the ingestion of ice crystals while operating at high altitude during a 1997 Honeywell flight test campaign investigating the turbofan engine ice crystal icing phenomena. The test plan included test points conducted at the known flight test campaign field event pressure altitude and at various pressure altitudes ranging from low to high throughout the engine operating envelope. The test article experienced a loss of power event at each of the altitudes tested.

Many studies have been performed to quantify the formation and evolution of roughness on ice shapes created in Appendix C icing conditions, which exhibits supercooled liquid droplets ranging from 1-50 µm. For example Anderson and Shin (1997), Anderson et al. (1998), and Shin (1994) represent early studies of ice roughness during short-duration icing events measured in the Icing Research Tunnel at the NASA Glenn Research Center. In the historical literature, image analysis techniques were employed to characterize the roughness. Using multiple images of the roughness elements, these studies of roughness focused on extracting parametric representations of ice roughness elements. While the image analysis approach enabled many insights into icing physics, recent improvements in laser scanning approaches have revolutionized the process of ice accretion shape characterization.

From November 2010 until May of 2011, NASA's Icing Remote Sensing System was positioned at Platteville, Colorado between the National Science Foundation's S-Pol radar and Colorado State University's CHILL radar (collectively known as FRONT, or ‘Front Range Observational Network Testbed’). This location was also underneath the flight-path of aircraft arriving and departing from Denver's International Airport, which allowed for comparison to pilot reports of in-flight icing. This work outlines how the NASA Icing Remote Sensing System's derived liquid water content and in-flight icing hazard profiles can be used to provide in-flight icing verification and validation during icing and non-icing scenarios with the purpose of comparing these times to profiles of polarized moment data from the two nearby research radars.

An experimental research effort was begun to develop a database of airplane aerodynamic characteristics with simulated ice accretion over a large range of incidence and sideslip angles. Wind-tunnel testing was performed at the NASA Langley 12-ft Low-Speed Wind Tunnel using a 3.5% scale model of the NASA Langley Generic Transport Model. Aerodynamic data were acquired from a six-component force and moment balance in static-model sweeps from α = -5 to 85 deg. and β = -45 to 45 deg. at a Reynolds number of 0.24x10⁶ and Mach number of 0.06. The 3.5% scale GTM was tested in both the clean configuration and with full-span artificial ice shapes attached to the leading edges of the wing, horizontal and vertical tail. Aerodynamic results for the clean airplane configuration compared favorably with similar experiments carried out on a 5.5% scale GTM.

Operational issues encountered by Apollo astronauts relating to lunar dust were catalogued, including material abrasion that resulted in scratches and wear on spacesuit components, ultimately impacting visibility, joint mobility and pressure retention. Standard methods are being developed to measure abrasive wear on candidate construction materials to be used for spacesuits, spacecraft, and robotics. Calibration tests were conducted using a standard diamond stylus scratch tip on the common spacecraft structure aluminum, Al 6061-T6. Custom tips were fabricated from terrestrial counterparts of lunar minerals for scratching Al 6061-T6 and comparing to standard diamond scratches. Considerations are offered for how to apply standards when selecting materials and developing dust mitigation strategies for lunar architecture elements.

We consider the heat transfer characteristics of an ideal concentric disk used in the Wiped-Film Rotating-Disk (WFRD) evaporator for the Vapor Phase Catalytic Ammonia Removal (VPCAR) water recovery system. A mathematical model is derived to predict the radial temperature distribution and its average over the surface of the disk as a function of system parameters. The model shows self-similarity of the temperature distribution and the existence of a dimensionless parameter S (ratio of heat flux to convection) that can be used as a criterion to optimize the thermal characteristics of the disk in order to approach uniform surface temperature. Comparison of the model to experimental data using global (infrared imager) and local (resistive temperature devices) measurements shows that agreement with the model depends on the ambient condition denoted by the local heat transfer coefficient.

Thermal characterization was performed on a vapor compression heat pump using a novel, hybrid two phase loop design. Previous work on this technology has demonstrated its ability to provide passive phase separation and flow control based on capillary action. This provides high quality vapor to the compressor without relying on gravity-based phase separation or other active devices. This paper describes the subsequent work done to characterize evaporator performance under various startup scenarios, tilt angles, and heat loads. The use of a thermal expansion valve as a method to regulate operation was investigated. The effect of past history of use on startup behavior was also studied. Testing under various tilt angles showed evaporator performance to be affected by both adverse and favorable tilts for the given compressor. And depending on the distribution of liquid in the system upon startup, markedly different performance can result for the same system settings and heat loads.

Smoke detection experiments were conducted in the Microgravity Science Glovebox (MSG) on the International Space Station (ISS) during Expedition 15 in an experiment entitled Smoke Aerosol Measurement Experiment (SAME). The preliminary results from these experiments are presented. In order to simulate detection of a prefire overheated-material event, samples of five different materials were heated to temperatures below the ignition point. The smoke generation conditions were controlled to provide repeatable sample surface temperatures and air flow conditions. The smoke properties were measured using particulate aerosol diagnostics that measure different moments of the size distribution. These statistics were combined to determine the count mean diameter which can be used to describe the overall smoke distribution.

NASA's Mars Science Laboratory mission will use a Powered Descent Vehicle to accurately and safely land a roving, robotic laboratory on the surface of Mars. The precision landing systems employed on this vehicle are exposed to a wide range of mission environments from deep space cruise to atmospheric descent and require a robust and adaptable thermal design. This paper discusses the overall thermal design philosophy of the MSL Powered Descent Vehicle and presents analysis of the active and passive elements comprising the Cruise, Entry, Descent, and Landing thermal control systems.