The Vertical Lift Research Center of Excellence Department of Aerospace Engineering The Pennsylvania State University, University Park, PA 16802

Transcription

1 ADVERSE ENVIRONMENT ROTOR TEST STAND CALIBRATION PROCEDURES AND ICE SHAPE CORRELATION Jose L. Palacios Research Associate Edward W. Brouwers Research Assistant Yiqiang Han Research Assistant Edward C. Smith Professor The Vertical Lift Research Center of Excellence Department of Aerospace Engineering The Pennsylvania State University, University Park, PA ABSTRACT An Adverse Environment Rotor Test Stand (AERTS) has been designed, and constructed. The facility is able to reproduce natural icing conditions on a hovering rotor. The motor/hub configuration is designed to spin instrumented rotors of up to 9 ft. diameter and has the capability of providing tip speeds of up to 470 ft/sec. A Liquid Water Content (LWC) calculation methodology was developed and sensitivity studies to determine experimental LWC are presented in this paper. Correlation between experimental ice accretion shapes obtained in the AERTS facility and experimental results obtained by the NASA Icing Research Tunnel and the Air Force Arnold Engineering Development Center are presented. These experimental correlations are conducted to demonstrate the capability of producing an accurate realistic icing cloud of the new facility. All tests reported in this paper have been conducted on 1 in. diameter circular cross section rotors. The majority of the experimental ice shapes compared agree with results presented in literature with thickness errors as low as 2% and impingement limits discrepancies no greater than 15%. Nomenclature Ac b cp,ws Accumulation parameter, dimensionless Relative heat factor, dimensionless Specific heat of water at the surface temperature, cal/g Khc Convective heat-transfer coefficient, cal/s m 2 K hg Gas-phase mass-transfer coefficient, g/s m 2 K Inertia parameter, dimensionless K0 Modified inertia parameter, dimensionless ka Thermal conductivity of air, cal/s m K LWC Cloud liquid-water content, g/m 3 Ma Mach Number, dimensionless MVD Water droplet median volume diameter, μm Nu Nusselt number, dimensionless m& Mass Flux, Kg/m 2 sec Pr Prandtl Number Pst Static pressure, psi pw Vapor pressure of water in atmosphere, psi pww Vapor pressure of water at the icing surface, psi r Recovery factor, dimensionless Re Reynolds number of model, dimensionless Reδ Reynolds number of water droplet, dimensionless τ Accretion time, min t Temperature, C tf Freezing temperature of water, C ts Surface temperature, C V Free-stream velocity of air, m/s β0 Collection efficiency at stagnation line, dimensionless δ Cylinder diameter, cm Ice thickness at stagnation line, cm ηa Freezing fraction (Messinger analysis), dimensionless ηe Freezing fraction, experimental, dimensionless θ Air energy transfer parameter, K λ Water droplet range, m Λf Latent heat of freezing of water, cal/g λstokes Water droplet range if Stokes Law applies, m Λv Latent heat of evaporation of water, cal/g μ Viscosity of air, g/m s ρa Air density, g/m 3 ρi Ice density, g/m 3 ρw Liquid water density, g/m 3 Ф Water droplet energy transfer parameter, K Presented at the American Helicopter Society 66 th Annual Forum, Phoenix, AZ, May 11-13, Copyright 2010 by the American Helicopter Society International, Inc. All rights reserved.

2 T 1. INTRODUCTION HERE are four main ways to perform rotor ice testing: wind tunnel ice testing, ground spraying approaches, in-flight spraying systems, and chasing the weather. Wind tunnel testing considerably reduces cost comparing with natural icing trials and it is ideal for ice shape model validation and preliminary testing of fix wing ice protection technologies. Wind tunnel icing facilities, such as NASA Glenn (OH) [1], Cox Icing Research Facility [2] (NY), or Goodrich Icing Tunnel (OH) [3] have limited test sections restricting the size of the rotor blades that can be spun. Ground icing facilities or in-flight water spraying system [4] (such as the Helicopter Icing Spray System, HISS) are expensive approaches to reproduce natural icing on full scale rotors and are not suitable for proof-ofconcept testing of new deicing systems. These are testing techniques used for certification of established ice protection systems. Similarly, chasing the weather involves high costs reserved for ice protection certification of technologies in production. A main obstacle on on-going efforts to model rotor ice accretion is lack of open source validation ice shape data. There are limited icing facilities focusing on rotorcraft research. Rotor icing testing can be accomplished in NASA Glenn s Icing Research Tunnel (IRT), but model rotor diameters are limited to 6 ft by the test section [1]. The Helicopter In-flight Spray System (HISS) allows for full scale testing, but detailed icing shapes are difficult to acquire as the vehicle must descent through layers of warm air that may shed ice prior to landing. Ice protection system and other components tests, such as icing sensors or effects of icing on other probes, can be conducted at above mentioned icing wind tunnels. These tests are often limited by tunnel velocity and the fact that the centrifugal forces inherent to rotor rotation are not represented. For this reason, a new Adverse Environment Rotor Test Stand facility (AERTS) has been designed, fabricated and calibrated. In this facility icing conditions can be reproduced surrounding a 9 ft. diameter rotor. The main mission of this facility is to provide a test bed for new ice protection systems, such as ultrasonic deicing. Secondary objectives involve measurement of ice adhesion strength to different coatings and ice shape correlations with ice accretion models. To determine if natural icing conditions can be reproduced, Liquid Water Concentration (LWC) must be properly characterized in the facility. This is one of the most important parameters used during ice accretion experimental testing [5] and it is measured in g/m 3. LWC sensors are not applicable to the facility because they require velocity over the active element. To provide these devices with proper operational velocity conditions, the LWC sensors would have to be spun. Due to size and cost of these sensors, their rotation was not possible. For this reason, in the AERTS facility, LWC is calculated using the modified accumulation parameter of a body, which involves the accurate determination of the freezing fraction. The freezing fraction of a body, η, measures the percentage of water that freezes to the shape with respect to all the water coming into contact with the body. Freezing fraction is dependent on icing conditions including LWC. Freezing fraction is therefore the most important non-dimensional parameter used in ice accretion modeling, scaling, and calibration as it represents all the effects from LWC, droplet size, temperature and other icing test parameters. The accurate determination of the freezing fraction of a body is critical to the calculation of LWC. This quantity is defined by Messinger s heat-balance analysis and it determines how rapidly freezing takes place when super-cooled water impacts a solid body. In this paper, a description of the facility is provided, as well as initial calibration procedures to determine LWC and MVD in the laboratory. To demonstrate the facility initial capabilities and to determine its application limits, ice accretion shapes to a 1 in. diameter cylinder are correlated to results presented in literature. 2. RESEARCH OBJECTIVE The objectives of the research are to describe the capabilities of the newly designed and constructed AERTS facility, the purpose of which is future evaluation of helicopter deicing systems. Analysis of LWC sensitivity studies is presented. To validate the functionality and capability to reproduce natural icing conditions on the rig, ice shapes obtained in the facility are correlated with prior testing conducted at the NASA Glenn Icing Research Tunnel [6] and the Air Force Arnold Engineering Development Center (AEDC) [11]. This paper is divided into three sections: - Facility Description - LWC Sensitivity Study - Ice Shape Correlation to NASA Icing Research Tunnel and AEDC Experimental Results on 1 in. Diameter Cylinder

3 3.1 Nozzle Spray System Array 3. AERTS FACILITY The AERTS facility is formed by an industrial 20 x 20 x 20 foot cold chamber where 4 in. thick insulated walls and a water-cooled compressor form the cooling system. Temperatures between -25 C and 0 C can be achieved in the chamber. The chamber floor is waterproofed with marine lumber covered by aluminum plating, and a drainage system in the perimeter of the room, collects melted ice during the post-test defrosting process. Inside the chamber, and surrounding the rotor, there is a safety ballistic wall in the shape of an octagon. The ballistic wall is formed by 6 in. thick weather resistant lumber reinforced with 0.25 in. thick steel, and covered by aluminum plating for weather protection. A schematic and photograph of the chamber, as seen from a top view, is shown in Figure 1. In the ceiling of the chamber there are 15 NASA standard icing nozzles that generate the icing cloud to the room. The nozzles are arranged into two concentric circles located 20 and 42 in. from the center of rotation...the nozzles can be operated in sets of five, having the capability to turn on five, ten or fifteen nozzles. The number of nozzles operating and the Median Volume Diameter (MVD) of the water droplets (provided by the pressure differential between air and water) dictate the Liquid Water Content (LWC) in the room. Similar nozzles are used in the Icing Research Tunnel (IRT) and Goodrich Icing Tunnel. A photograph of the icing cloud start is shown in Figure HP Motor Ballistic Wall 15 Nozzles Facility Ceiling View 20 Figure 2: Detail of Icing Cloud AERTS Facility 20 Cooling Fan Slip Ring Weather Station 9 ft Bell Housing w/ 6 Axis Load Cell Collective Actuator Figure 1: Schematic and Photograph of the AERTS Facility. The AERTS Hub is Collective and Lateral Cyclic Capable. Max RPM: Max. Rotor Diameter: 9 ft. MAX. Power: 120 HP. The nozzles operate by aerosolizing water droplets with a precise combination of water and air as per nozzle calibration curves[1]. The plots relating water and air pressure differential to the MVD particle size created is presented in Figure 3. The nozzles are installed in parallel, such that equal air and water pressure is sent to every nozzle. This is done because the pressures control the water particle size provided by the nozzles The air system was designed to provide accurate and consistent air pressure to the icing nozzles. Each nozzle requires up to 15 CFM to provide a stable icing cloud at 50 psi of input air pressure. A 21 HP air compressor provides the nozzles the required pressure, with an upper limit corresponding to constant 35 psi air pressure to all 15 nozzles, 55 psi to 10 nozzles, and 105 psi of air pressure to 5 nozzles. These upper limits dictate the airline pressures that can be triggered.

4 Air (Psi) MVD (Microns) P =Delta Pressure Water - Air (Psi) The icing cloud can be turned on and off at desired settings (MVD, airline) from the control room, where remote electronic shut-off valves of the water and air systems are located. Custom designed computer software controls the water and air pressures to desired settings using feedback control loops, maintaining the particle size within 2 µm of desired parameters. The remote capabilities of the cloud allow starting the icing once the rotor has reached desired RPM. Suction pumps are in place to stop the cloud instantaneously once the system is shut down. Figure 3: NASA Standard Icing Nozzle Operation Chart as Described in Reference 1 The water system is generally similar to the air system, with added complications in maintaining constant and controllable water pressure in a close loop. For this reason, a feedback control is in place to maintain the water pressure at desired conditions. In addition, a water reverse osmosis purification systems is required prior pressurization in order to reproduce natural icing conditions and to prevent nozzle clogging. The water purity measured 1 ppm and a resistance of 2 KΩ between two electrodes immersed in the water 6 in. from each other. The water and air pressures are measured at the input of the water and air lines to the nozzles, ensuring precise readings of the pressure differential controlling the particle size. A diagram of the air and water system is shown in Figure Motor/Hub In the center of the chamber a 125 HP, 160 ft-lb motor rotates the lower hub of a QH-50D DASH UAV vehicle. The motor is connected to a gear box with a 2.5:1 reduction ratio. The hub was retrofitted to fit the transmission of the motor. The configuration provides RPM values of up to 1500 RPM for 4.5 ft. radius blades, reproducing full scale helicopter tip speeds. The test stand has been successfully operated to date up to 1000 RPM. The hub has collective and lateral cyclic capabilities, as well as a six-axis load cell. A detail photo of the hub is presented in Figure 5 and a summary of key facility capabilities is listed in Table Controls and Measurements The facility is operated from a control room where remote controllers of all aspects of the facility are located. Controls are separated on three main independent groups: rotor, icing cloud and cooling. Figure 4: Diagram of Air and Water Conditioning Prior Atomization at Nozzle QH-50 Bell Housing/HUB Bell housing with Heated Load Cell Figure 5: Photo of QH-50D DASH UAV Hub

5 Parameter Value Table 1: AERTS Facility Capabilities Rotor System Notes Motor Power (HP) Motor Torque (in-lbf) 84 max Unlimited run time 120 max 3 minute run limit 696 max Unlimited run time 995 max 3 minute run limit RPM 400 to RPM for adequate cloud mixing; 1000 RPM test stand safety limit Rotor Tip Radius (ft) 2 to 4.5 Minimum required to reduce effects of hub icing cloud; maximum dictated by ballistic wall geometry Blade Grip Radius (ft) 0.94 Blade Grip CF Load (lbf) 14,000 Based upon QH-50 hub design loads. Includes required facility FOS Hub Precone ( ) 3 Hub Flap (Teeter) Range ( ) -12 to +12 Limited by teeter bumpers Collective Pitch ( ) -2 to +12 Controlled with linear actuator. Lateral Cyclic Pitch ( ) -5 to +5 Controlled with linear actuator. Longitudinal cyclic pitch is locked out. Parameter Value Active Nozzles 1 to 15 Temperature ( C) Ambient to -25 MVD (µm) 10 to 50 LWC (g/ m 3 ) 1 to 5 Icing Time (seconds) 30 to 240 Water Input Temp. ( C) 20 to 71 Water Purity (ppm impurities) 1 to 3 Icing System Notes Nozzles are arranged in 2 concentric rings in chamber ceiling, with 5 nozzles in inner ring and 10 nozzles in outer ring. Any combination of nozzles can be used for each test. Chamber cooling system is shut down during each test to avoid disrupting icing cloud. Temperature increases during each test due to kinetic friction of the rotor and warm water inputs to the chamber. Not directly measured. Droplet size based upon NASA Standard nozzle calibration tables. Water and air pressure input control system maintains droplet size input to the test chamber ± 2 µm. Larger particle sizes are possible, but are outside the calibrated range of the nozzles. Not directly measured. Controlled by number of active nozzles and input pressures and calculated after each test based upon accreted ice thickness. Not all MVD/LWC points are possible based upon chamber limitations. Approximately 30 seconds is required for the nozzles to stabilize and for the cloud properly to mix in the chamber. Icing duration limit is based on the requirement to maintain static temperature in chamber ± 1 C from desired point. Water input temperature can be varied to properly supercool droplets at all temperatures. Reverse Osmosis Purification System generates water with 2 MΩ resistance (6" probe separation) Parameter Value Instrumentation Notes CCD Cameras 3 Two cameras monitor entire test chamber; third camera focused on rotor tip. Static Temperature Sensors 5 Slip Ring Power Channels 24 max Each channel is rated up to 15A, 800V Slip Ring Signal Channels 24 max Each channel is rated up to 2A, 100V Thermistors are positioned around test chamber to monitor internal temperature. A commercial weather station measures temperature as well as pressure, humidity and rain fall rates.

6 6 Axis Load Cell Forces (lbf) 6 Axis Load Cell Moments (in-lbf) Thrust: 875 Lat/Long: 300 Pitch/Roll/Yaw: 1800 Shaft Torque Sensor (in-lbf) 1500 Measurement range. Available Output Channels 8 Available Generic Input Channels Available Strain Gauge Modules 38 (+4) Strain Gauge Amplifiers 2 Accelerometer Conditioning Circuits 2 Measurement range. Sensor can safely handle 8x reported loads. Each channel is rated at 0-10V, 10 ksps. Current is based upon DAQ module usages, but typically ~ 10 ma. Used for triggering relays etc. Each channel is rated at 250 ksps. Four channels are reserved for temperature (thermocouple) measurements. 3 Used only for fixed frame measurements Strain gauge amplifiers are used for on blade measurements. The full bridge completion units are mounted at the blade roots and amplify signals prior to them entering the rotor hub to improve overall S/N ratio. Accelerometer signals are conditioned and amplified prior to being read by Labview The rotor system has ramping capabilities, currently set-up to reach 1000 RPM in 5 seconds. Emergency stop of the rotor allows for complete stop in 2 seconds. The walk-in freezer is cooled by convection of cooling lines and a fan located inside the chamber. To prevent the fan from accreting ice and distorting the icing cloud, the fan is turned off during exposure to supercooled water droplets. This currently limits the capability to maintain a desired temperature within 1 C to 3.5 minutes, as warm air and water and kinetic friction of the rotor increase the temperature in the chamber. A six axis load cell is installed in the rotor stand. It measures lift, side forces and related moments. This load cell does not only provide physical loads on the system, but also monitors for potential rotor unbalance due to ice shedding. To monitor ice accretion performance degradation, a torque sensor placed in line with the shaft measures torque on the system with a maximum reading capability of 95 ft-lbs. Accelerometers are placed in the bell-housing and mast as redundant measures to monitor potential vibration due to rotor unbalance. Inside the bell-housing and monitoring the rotation of the main shaft, there is an RPM sensor that provides information on the rotor status. Temperature in the chamber is monitored by weather stations inside of the ballistic wall and under the rotor plane. Additional thermocouples are placed on the room, also for redundancy. All electric systems and signal conditioners are insulated and heated to 29 C to limit temperature compensation errors on the readings. 4. LWC SENSITIVITY STUDY S Static temperature is measured in the facility during icing testing, as it can be read from thermocouples located in the chamber. MVD is calculated from NASA calibration tables and experimental readings of pressure differentials between the water and the air inputs to the nozzles. Even though MVD is not currently directly or indirectly measured in the facility, however, nozzle calibration is assumed to be accurate and constant monitoring of the water and air pressure allows for the calculation of MVD. The only icing parameter that is unknown and could not be directly measured during testing is the LWC in the room. LWC sensors, in addition to their high cost, require velocity components that are not available in the facility. These sensors are designed for wind tunnel and in-flight LWC measurements and require a minimum velocity component of 15 m/s. To subject the LWC sensors to these velocities, they could be placed in the rotating frame, but due to the size of these sensors and centrifugal effects, their rotation is not possible. For this reason, LWC must be determined experimentally by investigating the ice accretion thicknesses to a known body. These calibration efforts must be performed prior attempting ice shape correlations with literature or ice accretion model validation. A computer code that calculates LWC from ice accretion thickness during a given time interval was created. One of the most important non-dimensional

7 parameters used to calculate LWC modeling is the freezing fraction, η, defined by the heat-balance analysis of Messinger. The code was validated versus experimental and analytical results presented by Anderson and Tsao [7] on their paper Evaluation and Validation of the Messinger Freezing Fraction. To determine the LWC in the facility, a computer code that calculates LWC from ice accretion thickness during a given time interval was created. The principle of this code traces back LWC from the experimental result of ice accretion thickness. The code correlates thickness and freezing fraction to determine the experimental LWC. The freezing fraction is defined by Messinger [8] as the fraction of water flux entering a control volume that freezes within the control volume. It illustrates the ice accretion rate when super-cooled water impinges on a solid body, on which the ice thickness is depended. In this way, the LWC can be calculated from the experimental ice thickness. This code was validated versus experimental and analytical results from reference 7. The calculation scheme is represented in the following sections. 4.1 LWC Experimental Calculation The developed code estimates the LWC from the stagnation point ice thickness for a given accretion time. The estimated LWC, together with other icing conditions, is then compared with analytical results presented in literature [7] to validate the procedure. The input parameters to the code include: chord (for airfoil) or diameter (for cylinder), MVD, temperature, local velocity, icing time and thickness. To calculate the physical LWC condition during testing with this analytical method, several parameters are introduced in the analysis and are described in the following sections Droplet Trajectory Analysis Prior Impingement The objective for this section of the analytical calculations is to find the collection efficiency, which can be interpreted into how much water droplets are going to hit on the model (i.e. the mass flux used in the following equations). This is the basis of both analytical and experimental expression of freezing fraction. This analysis aims to find an expression for supercooled water drop distribution. The stagnation line collection efficiency, βo, illustrates the impinging water drop trajectory by considering the projection of a stream tube from the far-field inflow at stagnation line. The problem is simplified at the stagnation line, as it is assumed that at this line there is no incoming interference from other controlled volumes. The analysis following are all based on this assumption [7,8]. The expression of collection efficiency at the stagnation line is given by Equation 1:.84 ( K ) ( ) 0 1/ 8. K0 1/ * β 0 = 84 (1) * where, K0 is the Langmuir and Blodgett s [9] expression for modified inertia parameter (Equation 2). This equation was initially published for cylinders, but was then validated for airfoils in the reference 9. 1 λ 1 1 K 0 = + K, for K > (2) 8 λ Stokes 8 8 The inertia parameter, K, in Equation 2 can be expressed as: 2 ρwδ V K = (3) 18dµ And Stokes a λ / λ is defined as the dimensionless droplet range parameter, λ 1 = λ Re Re Stokes where Vδρ δ δ (4) Re = a δ (5) µ a Energy Balance Analysis during Impingement As mentioned, one of the most important variables during icing testing is the freezing fraction, which denotes the fraction of water droplets that freezes at the surface of a body, thus indicating the heat balance at the ice surface. Analytical freezing fraction can be found by the following Equation: C p, ws θ n0, a = φ + (6) Λ f b where, ф and θ, are defined as droplet energy transfer and air energy transfer coefficients respectively: 2 V φ = t f tst (7) 2c V h p, ws p p 2 G ww w θ = ts tst + Λ v c p a h c p (8) 2, st The relative heat factor, b, is introduced by Tribus [9] as:

8 mc & p, ws b = (9) h c The convective heat-transfer coefficient, hc, can be calculated from Equation 11. hcd Nu = (10) k a The numerical expression of Nu in this code is chosen according to different Re numbers: for Re > 10 5, as per reference 9: Nu = 1.10Re (11) and for Re < 10 5, as per reference 9: Nu = 114. P r Re (12) Based on the trajectory analysis at stagnation line in the last section and assuming βo and ρi remain the same while the ice shape changes during the test, the mass flux can be expressed as: m& = LWC V (13) Here, it can be seen that LWC can be determined from the analytical freezing fraction. by introducing a correlation between freezing fraction and ice thickness in next section, the LWC can be finally determined Ice Accretion Analysis Based on the previous analysis, a time-span analysis during ice accretion can be performed. Total ice thickness at stagnation line,, can be expressed as: m& τ = (14) ρ i n 0 By Substituting Equation 10 into Equation 14 and introducing an accumulation parameter Ac, Equation 15 is found. LWC V τ Ac = (15) d ρ i The non-dimensional total ice thickness is defined in Equation 16. 0, e A c β0 d = n (16) The experimental freezing fraction, η0,e, can be related to the analytical freezing fraction, η0,a, by using a linear curve fitting as it is suggested by Anderson and Tsao [7] : n 0, e = n0, a (18) β 0 The relationship between total thickness and LWC can be shown to be monotonic. Thus, an exhaust algorithm can be implemented to find experimental LWC from total ice thickness per time. The scheme of the code is summarized in Figure Evaluation of LWC Calculation Code The calculated LWCs based on the total ice thickness per time are compared with the analytic LWCs presented in literature for both cylinders [10] and airfoils [7]. The correlation between calculated LWC and results presented in literature are shown in Figure 7 and Figure 8. It can be concluded that this code calculates acceptable LWC from total ice thickness per time (within ±15% error) for nearly 90% of all the cases presented in literature. Taking into account the uncertainties related to experimental test data, these results can be assumed to be useful and reliable to support the LWC calibration of the facility Uncertainty Analysis From Figure 7, and Figure 8, it is shown that experiment-derived LWCs generally result in a good agreement with literature data, presenting correlation discrepancies of less than 15% for the majority of the cases compared. Several cases deviate between calculations and experimental results presented by the referenced documents. The two main contributions of this kind of error come from uncertainty of measurement; and error transmitted between calculation equations. Firstly, for most experiments performed at NASA IRT to which this paper is comparing, the uncertainty related to LWC calibration at IRT is claimed to be about ±12% [7]. Also there is ±12% uncertainty in MVD. ] In addition, in most icing tests, hand-tracing measurement methods are prevalently used, and for this reason, the thickness record has its own inherent uncertainty. For similar shapes, it can be shown that the experimental ice thicknesses can differ by up to 18.8% [7] between the centerline of a test section and some small distance above centerline. Given the limited data set, these uncertainties cannot be effectively resolved.

9 Figure 6: Scheme of Experimental LWC Calculation Code Figure 7: Cylinder LWC Calculations from Total Thickness and Correlation with Results Presented in Reference 10 Figure 8: NACA 0012 LWC Calculations Compared to Experimental Results Presented in Reference 7

10 Secondly, due to the small size of the ice thickness itself, a slight error in tracing the ice thickness will then be transmitted and amplified through equations and computing loops of the presented code, resulting in a relatively big error between analytic LWC and thicknessbased or experimental LWC. It can be seen from presented equations, that the change in thickness has large effects on the calculated LWC. As stated before, there is a linear relationship between thickness, freezing fraction and eventually the LWC. Small changes in ice thickness (> 0.5 mm) will produce deviations of LWC of up to 50%. In the reference [7], although with a different analysis method and ignoring the difference between analytical and experimental LWC, Anderson and Tsao also did some comparisons between analytical freezing fraction and experimental freezing fraction based on the ice thickness. In two test groups (test case number 8 14 and with regard to Figure 8 in this article), large discrepancies between ηa and ηa can be found in these two groups. The greatest one is found in case /1(test case number 32 in Figure 8, with regard to this article), where ηa = and ηe = 0.190; i.e., the error can be as high as 45% (error with respect to ηe, from which the experimental LWC is determined), much bigger than ±12% as they expected for most cases. These errors are also reflected in the LWC calculation code in Figure 8. The same phenomena are also found in Figure 7, test case number 9 and 10. Anderson and Tsao believe this is because there can be significant uncertainty in the ice thickness values found from tracings at low freezing fractions. This is true as already mentioned above. Also, the relatively large discrepancies between the analytical LWCs and the ones calculated from the measured thickness can also be explained by the slope of the relationship between ice thickness and LWC (Equation 19). The slope, S, of the equation could be very small ( 0.025), greatly affecting the LWC value for errors introduced in the measurement of the ice thickness. (19) = S LWC + r For example, in some cases, a change in thickness of in. results in a change on the calculated LWC of 0.43 g/m3. For this reason, careful measurement of the ice shapes must be performed. In addition, the empirical equations used in this code (such as relationships between ηe and ηa, or the numerical expression of Nu.) will add error into the calculation as icing conditions diverge from those used during the definition of these empirical equations. With these assumptions of uncertainty, each analytic LWC is plotted in Figure 7 and Figure 8 with an error bar of ±15%. Calculated LWC results correlate with values presented in literature, validating the usage of the code to determine the LWC in the facility. 4.2 Experimental Results A rotor formed by a 1 in. diameter cylinder (50 in. radius) was spun at different icing conditions. A total of 18 runs were conducted to calibrate the chamber at -5 C, -10 C, and -15 C. A 1 in. diameter rotor of 50 in. radius was used. For each temperature, variations of RPM (500, and 600), air lines (20, 25, and 30 Psi) and MVD (20, 25, and 30 µm) were conducted. LWC were calculated for all conditions along the span of the rotor, as ice thickness. All tests were run for 3 minutes, with a maximum temperature deviation of 1 C. Five nozzles located in the outer ring were used during testing. LWC calibration matrices allow for the selection of conditions (MVD, airline, temperature) to trigger a desired experimental LWC during a test. A photograph of the 1 in. diameter rotor as seen from one of the monitoring cameras during ice testing is presented in Figure 9. 1 in. Diameter 50 in. Bottom View Rotor with Accreted Ice Figure 9: Photograph 1 in. Diameter Rotor during Icing Testing As it can be observed on Figure 10, the ice thickness at the stagnation point increased as temperature decreased from -5 C to -15 C. These tests were conducted at ceteris paribus conditions (25 MVD, 500 RPM, 25 Psi air line, 3 minutes of ice exposure). It is counterintuitive that the ice thickness would increase as temperature drops, as the LWC should decrease. This does not happen in the AERTS facility for all conditions, as temperatures below -10 C are allowing more droplets to become super-cooled. For this

11 reason, each temperature condition at the facility must be calibrated for LWC. A detail of the ice thickness increase is depicted in Figure 11, where the calculated LWC is shown. Ice Thickness (in.) C 500 RPM, 25 Psi Air, 25 MVD, 3 min C C Rotor Span (r/r) -15 C -10 C -5 C Figure 10: Variation on Ice Thickness with Rotor Span: -5º, -10º and -15º Deg. C Static Temperature RPM increases correspond to ice increases, as it is shown in Figure 12. This was expected, given the reduced tip speeds (about 65 m/sec), in where kinetic heating of the blade is not a major factor affecting ice accretion. Also, as particle size was increased between 20 and 30 MVD, ice accretion thickness increased C, LWC, 2.55 g/m C, LWC 2.35 g/m C, LWC 2.29 g/m 3 decrease the maximum air pressure line to be used. To allow for the use of higher airlines, a decrease in operational nozzles could be implemented. Ice Thickness (in.) Figure 12: Effect of RPM Increase on Ice Thickness and LWC One important issue encountered during testing was the appearance of ice crystals when air lines exceeding 23 Psi (see Figure 3) were triggered. Due to the small facility size, super-cooled liquid droplets re-circulate around the ballistic wall after they pass through the rotor plane if they do not accrete to the walls or floor of the facility. Since no particle removal process is used in the facility (other than ice collection screens located under the rotor plane) the droplets can freeze into solid crystals when they re-circulate. Ice Thickness (in.) RPM 600 RPM 500 RPM, 25 Psi Air, -15 Deg. C 20 MVD 30 MVD Linear (30 MVD) Linear (20 MVD) LWC = 2.4 g/m 3 25 Air line, 25 MVD, -5 Deg C. LWC 2.4 g/m 3 LWC 3 g/m Span Location (r/r) LWC = 2.2 g/m Rotor Span (r/r) Figure 13: Effect of MVD Increase on Ice Thickness and LWC 25 Psi Air, 25 MVD, 500 RPM 30 Psi Air, 25 MVD, 500 RPM Figure 11: Detail of Ice Shapes Obtained at -15 C, -10 C, and -5 C (500 RPM, 25 MVD, 25 PSI Air Line, 3 min. Exposure) For this reason, the facility is limited to the 30 Psi air line if 25 MVD are sought and 5 nozzles are in operation. This airline limitation will vary depending on the MVD sought, as this is controlled by the pressure differential between water and air. Lower MVD will allow for an increase in the airline, while larger MVD will further Ice shape Erosion Figure 14: Effect of Airline Increase (-10 C). Notice Ice shape Erosion, as Facility is Saturated by Ice Crystals 30 Psi

12 When liquid droplets impact a crystal, the droplet is immediately crystallized, which creates a chain reaction [5]. Larger numbers of particles in the chamber increase this effect due to saturation and are generated when using higher air pressure inputs to the nozzle. To maintain a desired MVD at higher air pressures, water pressures need to be increased to maintain the proper pressure differential, as detailed in Figure 3 and explained in Reference 1. Since the water flow rate is dependent on this pressure differential, the mass of water added to the chamber increases, creating a large number of droplets. If the droplets crystallize, they erode ice shapes, providing spear shaped ice accretion, as shown in Figure 14. The maximum pressure differential to avoid crystallization problems was experimentally determined to be 23 psi. Similar ice shape erosion is documented by Tsao et al. in reference 12. During tests conducted at the IRT, there was evidence indicating that ice erosion occurred for rime ice shapes obtained at 250 knots. Erosion was identified by shapes lacking expected small-scale feathers and increased stagnation ice thickness, as seen in Figure 15 [12]. fully populated yet, as this is the first attempt to understand and calibrate the facility. For this reason, specific conditions identified in literature are cumbersome to match perfectly in terms of LWC. Despite these temporary limitations, agreement between ice shapes presented in literature and experimental results are observed. Correlations between AERTS experimental results and experimental results presented in literature (Reference 6) are shown in Figure 16 and Figure 17. As it can be seen in Figure 16 and Figure 17, the stagnation ice thickness correlates with experimental results obtained at the IRT. The overall shape of the accreted ice also agrees. Increases in the impingement limits can be observed, which could imply an increase in particle size or decrease in temperature during testing. AERTS EXP. 25 MVD 510 RPM r/r 58 m/sec C Ice shape Erosion Tsao, J., Kreeger, R., Reference 12 Figure 15: Example of Eroded Rime Ice Tracing at the IRT, Reference AERTS ICE SHAPE CORRELATION C TO NASA AND A AIRFORCE EXPERIMENTAL RESULTS To validate the capability of the facility to reproduce natural icing conditions, accretion shapes found in literature for 1 in. diameter cylinders were compared to experimental results obtained in the AERTS facility. Currently, the main challenges operating the facility are due to lack of temperature control and complete LWC calibration. Radiation cooling systems capable of maintaining the facility at temperature without convection fans are not installed yet. The temperature in the facility can only be maintained for 3.5 minutes with a deviation of 1 C, since all cooling fans must be shut down during icing to avoid ice accretion and cloud perturbation. A second issue is that LWC calibrations matrices are not Figure 16: Test 1 - Correlation of Experimental Results from AERTS (25 MVD, 58 m/sec C, 5 min., 2 gr/m 3 ) to Reference Results (Ref. 6: 23 MVD, 58 m/sec, C, 5 min., 1.6 gr/m 3 ) The discrepancy between impingement limits of the two experimental results was calculated to be less than 16% of the total ice thickness for both Tests. AERTS Exp. Reference AERTS Exp. Reference AERTS EXP. 1 in Tube 27 MVD 490 RPM 0.91 r/r 59.2 m/sec C Figure 17: Test 2 - Correlation of Experimental Results from AERTS (27 MVD, 59.2 m/sec C, 6.3 min., 1.91 gr/m 3 ) to Reference Results (Ref. 6: 27 MVD, 58 m/sec, C, 6.3 min., and 1.3 gr/m 3 )

13 Correlations were also performed against experimental ice shapes obtained by Ruff et al. at the Air Force Arnold Engineering Development Center (AEDC) [15]. These tests were performed at lower LWC values than the minimum provided by the facility when 5 NASA standard nozzles are in operation. For this reason, and to reduce the LWC in the chamber, a controlled system of nozzles was introduced, such that each nozzle can be operated individually. This allowed for combinations of 3 and 4 nozzles to be used, thus allowing for a reduction of LWC. As it can be observed on Figure 18 to Figure 22, ice shape agreement is obtained between all tests. MVD = 20 µm TRef = -15 C TExp = -15 C Vel = 60.9 m/sec LWCref = 1.2 gr/m 3 LWCExp = 1.3 gr/m 3 r/r = 0.91 Time: 3.75 Min MVD = 20 µm TRef = C TExp = -12 C Vel = 60.9 m/sec LWCref = 0.9 gr/m 3 LWCExp = 0.8 gr/m 3 r/r = 0.91 AERTS Exp. Reference Figure 20: Test 5 - Correlation of Experimental Results from AERTS to Reference Results Presented in Literature (Ref. 11) MVD = 20 µm TRef = -5 C TExp = -5.5 C Vel = 60.9 m/sec LWCref = 1.2 gr/m 3 LWCExp = 1.3 gr/m 3 r/r = 0.91 Time: 2.5 Min AERTS Exp. Reference Time: 2.5 Min AERTS Exp. Reference Figure 18: Test 3 - Correlation of Experimental Results from AERTS to Reference Results Presented in Literature (Ref. 11) The maximum discrepancy between ice thicknesses is calculated to be 11.8% for test 4 (Figure 19). Tests 1, 2, 5, and 6 have an ice thickness discrepancy between facilities of less than 2%. Deviations between targeted temperature and experimental temperature, in addition to other experimental uncertainty during tests, might introduce errors observed between shapes. In general, ice shape trends agree between experimental results presented in literature and results obtained at the AERTS facility, validating its capabilities to reproduce icing conditions. Time: 5 Min MVD = 20 µm TRef = -15 C TExp = C Vel = 60.9 m/sec LWCref = 1.2 gr/m 3 LWCExp = 1.2 gr/m 3 r/r = 0.91 AERTS Exp. Reference Figure 19: Test 4 - Correlation of Experimental Results from AERTS to Reference Results Presented in Literature (Ref. 11) Figure 21: Test 6 - Correlation of Experimental Results from AERTS to Reference Results Presented in Literature (Ref. 11) In the first test presented in Figure 16, the experimental MVD was 25 μm, 2 μm larger than what is presented in the result presented in literature. In the second test (Figure 17), the MVD was maintained at exactly 27 μm, matching the MVD presented in literature results. Discrepancies in MVD between the experimental tests compared are not believed to be the main cause of ice shape deviation. As mentioned on section 3, the MVD in the facility is maintained with a feedback control loop that ensures desired air and water pressure to the NASA standard nozzles. According to NASA nozzle calibration tables (Figure 3), the particle size is maintained within 2 μm. For this reason, MVD deviation is assumed not to be the main cause presenting the slight differences between both tests. It is believed that the main reason making the ice shape correlation deviate is that the temperature in the chamber can only be maintained for 3.5 minutes, before temperature increases exceed 1 C. Those tests run for longer than 3.5 minutes show increased discrepancies between experimental results. Tests 1 and 2 (Figure 16 and Figure 17) were run for 5 and 6.3 minutes respectively, having temperature increases of up to 2 C with respect to the desired starting temperature. From those tests

14 compared to AEDC results, tests 4, 5 and 7 were also run for more than 3.5 minutes, presenting larger shape deviations than tests 3 and 6 (run for 2.5 minutes). Time: 5 Min Figure 22: Test 7 - Correlation of Experimental Results from AERTS to Reference Results Presented in Literature (Ref. 11) This warming effect is believed to be the main source of any discrepancy that might be found between the two experimental results. Other uncertainties would be introduced due to the fact that the AERTS facility introduces centrifugal forces not seen in the IRT. 6. CONCLUSION MVD = 20 µm TRef = -5 C TExp = -6.7 C Vel = 60.9 m/sec LWCref = 1.2 gr/m 3 LWCExp = 1.32 gr/m 3 r/r = 0.91 AERTS Exp. Reference A new Adverse Environment Rotor Test Stand facility in were icing clouds surrounding a hovering rotor can be reproduced, was designed, and built to investigate ice accretion phenomenon and solutions. The AERTS facility is capable of reproducing natural icing conditions as long as saturation of the chamber is prevented. LWC sensitivity study of the facility for particle sizes between 20 and 30 MVD was accomplished. These efforts demonstrated that representative LWC values encountered in natural icing conditions (1.7 to 2.6 g/m 3 ) can be reproduced. From these initial calibration efforts, the saturation limit of the chamber was determined. This condition is identified when pressure differentials in the facility exceeded 23 Psi for a five nozzle configuration. During LWC sensitivity tests, it was also noted that the facility is limited to icing tests of less than 3.5 minutes due to temperature increases in the chamber. This temperature increases are due to lack of cooling systems during operation since they must be shut down to avoid ice accretion to cooling fans. This issue could be mitigated with additional radiation cooling lines in the facility that would allow for temperature control for longer periods of time. Ice shape correlations between the facility and experimental results presented in literature by NASA and Airforce, indicate the capability of the AERTS facility to reproduce icing shapes obtained in the IRT and the Arnold Engineering Development Center (AEDC). Correlations between IRT and AERTS stagnation ice thicknesses are excellent, with less than 2% discrepancy between tests. Impingement limits and overall ice mass was overachieved at the AERTS facility by up to 16%, due to experimental uncertainties, but it is believed that increases of temperature during testing beyond the desired comparison value are the main cause. These errors were calculated as percentage of thickness at the stagnation line. The maximum ice thickness errors with respect to the AEDC was calculated to be 11.8%, but the majority of tests provided correlations between ice thickness with discrepancies of less than 2%. Ice shapes obtained at the AERTS facility agree with experimental results presented in literature, validating the capability of the facility to reproduce natural icing conditions on hovering rotors with zero thrust. ACKNOWLEDGEMENTS The authors would like to thank Eric Kreeger and Paul Tsao of the NASA Glenn Research Center for their donation of the critical icing nozzles and their advice in calibrating the AERTS Facility. The authors would also like to acknowledge Peter Papadakos of the Gyrodyne Historical Foundation for the donation of the QH-50D lower rotor head and upper controls. The authors would also like to thank the US Army for sponsoring this research by funding a program to investigate the potential of the facility to generate representative ice shapes, and the installation of the icing system. This research is partially funded by the Government under Agreement No. W911W The U.S. Government is authorized to reproduce and distribute reprints notwithstanding any copyright notation thereon. The views and conclusions contained in this document are those of the authors and should not be interpreted as representing the official policies, either expressed or implied, of the U.S. Government. REFERENCES 1. Ide, R., Oldenburg, J., Icing Cloud Calibration of the NASA Glenn Icing Research Tunnel, AIAA , March Al-Khalil, K., Salamon, L., Tenison, G., Development of the Cox Icing Research Facility, 36 th Aerospace Sciences Meeting & Exhibit, AIAA , January 12-15, 1998, Reno, NV

15 3. Herman, E., Goodrich Icing Wind Tunnel Overview, Improvements and Capabilities, AIAA , 44 th AIAA Aerospace Sciences Meeting and Exhibit, 9-12 January 2006, Reno, Nevada 4. Peterson, A., Oldenburg, J., Spray Nozzle Investigation for the Improved Helicopter Icing Spray system (IHISS), 28 th Aerospace Sciences Meeting,, January 8-11, 1990, Reno, NV 5. Gent, R.W., Dart, N.P and Candsdale, J.T. Aircraft Icing. Philosophical Transactions of the Royal Society of London SeriesA. 2000, Vol Anderson, D., Rime-, Mixed-, and Glaze-Ice Evaluations of Three Scaling Laws, NASA Technical Memorandum , AIAA , AIAA 32 nd Aerospace Sciences Meeting and Exhibit, Reno, Nevada January 10-13, Anderson D., and Tsao, J., Evaluation and Validation of the Messinger Freezing Fraction, NASA/CR , AIAA , 41 st Aerospace Sciences Meeting and Exhibit, Reno, Nevada, January 6 9, Messinger, B.L., Equilibrium Temperature of an Unheated Icing Surface as a Function of Airspeed, J. Aeron. Sci. vol. 20 no. 1, January 1953, pp Anderson, David N., Manual of Scaling Methods, NASA CR Anderson, David N., Evaluation of Constant- Weber-Number Scaling for Icing Tests, AIAA and NASA TM , January Ruff, G., Analysis and Verification of the Icing Scaling Equations, Air force Technical Report AEDC- TR-85-30, November Tsao. J., Kreeger, R., Experimental Evaluation of Stagnation Point Collection Efficiency of the NACA 0012 Swept Wing Tip, AIAA , 1st AIAA Atmospheric and Space Environments Conference, June 2009, San Antonio, Texas