1 American Institute of Aeronautics and Astronautics

T-Range: The Navy�s Sea-Level Engine and Aerothermal Test Facility

W.K. Jaul* and S.L. Fitzpatrick� Naval Air Warfare Center Weapons Division, China Lake, CA 93555

In this era of defense downsizing and budgetary constraints, missile programs needing air-breathing engine or aerothermodynamic (aerothermal) testing must scrutinize their mission requirements and find test venues that can provide useful data in a cost effective and timely manner. The NAWCWD ground test facility, T-Range, has filled a niche in the Department of Defense (DoD) by providing low altitude supersonic conditions for both air-breathing engine and aerothermal testing. Ground testing of missiles and missile components can be a cost-effective means of risk reduction and down-selection if it accurately simulates the actual free-flight environments. This paper discusses the operational capabilities of T-Range and the method by which flight profile matching is performed. Programs that have used T-Range include: Evolved Sea Sparrow Missile (ESSM), Rolling Airframe Missile (RAM), Low-Drag Ramjet, Fasthawk, the Army�s Line of Sight Anti-tank (LOSAT) and Compact Kinetic Energy Missile (CKEM), and Ducted Rocket Engine Program. This paper also discusses the improvements that T-Range is implementing, including a new higher temperature air heater and supersonic nozzle, that will be available in early 2005 and will ensure that T-Range can meet the future testing needs of the DoD community.

I. Introduction Prior to the first flight of a new or improved Missile system designers often require that materials and

components first experience flight-like conditions of heating and shear in a ground test facility. Characterizing the performance of new airbreathing propulsion systems also relies on ground tests. In the ground test the missile component or combustor sees a high temperature stream of gas, just like in flight. Ground testing of missiles and missile components can be a cost-effective means of risk reduction and down-selection if it accurately simulates the actual free-flight environments. The NAWCWD T-Range, located at Naval Air Weapons Station (NAWS) at China Lake, California is a blowdown-to-atmosphere hot-gas supersonic test facility. Programs that have used T-Range include: Evolved Sea Sparrow Missile (ESSM), Rolling Airframe Missile (RAM), Low-Drag Ramjet, Fasthawk, the Army�s Line of Sight Anti-tank (LOSAT) and Compact Kinetic Energy Missile (CKEM), and Ducted Rocket Engine Program.

II. Description of Facility Testing Capabilities

A. Hot Gas Flowrate Capabilities T-Range is located outdoors at an elevation of approximately 2300 feet above sea level. High-pressure air is

stored at up to 3000 psi. Air storage capacity is currently 2900 cubic feet, providing about 47,000 pounds mass (lbm.) of pressurized air for testing. (Note: Approximately 50 to 60 percent of the stored air can be used before blowdown is affected due to low pressure.) Maximum test duration ranges from 2 minutes at 200 lbm/s, to 150 minutes at 3 lbm/s. The air is preheated in a sudden expansion (SUE) heater fueled by propane, where the combustion products are mixed with delivered air to increase its total temperature. For freejet testing, the hot gas is

* Mechanical Engineer, Aeromechanics and Thermal Analysis Branch, 1900 N. Knox Rd., Code 476100D, China Lake, CA 93555, warren.jaul@navy.mil. � Aerospace Engineer, Propulsion Research Branch, 1900 N. Knox Rd., M/S 6204, China Lake, CA 93555, shannon.fitzpatrick@navy.mil, AIAA member.

40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit11 - 14 July 2004, Fort Lauderdale, Florida

AIAA 2004-3658

This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

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then expanded through a supersonic nozzle to increase its velocity before it is exhausted over the test object. T-Range is currently capable of delivering air with a total temperature that can be varied from ambient to 2200°F at a total pressure of 1500 down to 16 psia. Heated air mass flow rates vary from 3 to 250 lbm/s. Heat fluxes as high as 300 Btu/ft2/s at the stagnation point of a 2-inch diameter hemisphere are possible for material testing.

The facility can be used to realistically simulate both the captive carry and free-flight portions of a missile�s flight profile. An explosion-fortified building, with monitoring systems for controlling the tests, houses the data acquisition system. High-speed video and film cameras (up to 8000 fps) and closed-circuit TV systems are used as needed for monitoring the tests. Pyrometers and two-color IR imaging cameras are available to measure surface temperatures over a wide test area. This has been extremely valuable when testing ablative materials.

While some aerothermal facilities cannot tolerate hardware failure and the resulting debris being blown downstream of the test location, T-Range can handle tests where component failure is likely to occur. These include ablative radomes, IR domes, and inlet port cover ejection tests. Tests at T-Range can include explosive devices for shroud removal. Up to 250 lbm of Class 1.1 energetic material can be used at T-Range and up to 2000 gallons of liquid fuel can be tested.

The T-Range facility has various nozzles with exit Mach numbers that include 1.0, 1.32, 1.92, 2.5 and 2.97. With its ability to generate air at 2200 °F, the Mach 2.97 nozzle can be used to simulate the heating environment of sustained Mach 4 flight at sea level with an exit diameter of 9.2 inches. The current water system can provide up to 2000 gallons of water at 2000 psi for cooling customer hardware.

Two test cells exist; cell I is used primarily for air-breathing engine testing and has a 50,000 lbf thrust stand to accomodate test hardware. Cell II is used for aerothermal testing and uses a hydraulically actuated ram to insert test items into the hot gas stream.

A schematic of the facility is shown in Figure 1. Air is transported from the 3000 psi storage vessels through a 6-inch schedule 160 stainless steel pipe. Airflow to each of the test stands is controlled by a 8-inch Hamel-Dahl valve operated by an electro-hydraulic actuator. A NIST traceable calibrated venturi is used to measure the air mass flow. The airflow rate is calculated using the measured total pressure and temperature at the entrance to the venturi and the area of the venturi throat. Since air does not behave as a perfect gas over the temperature and pressure ranges encountered during testing, a critical flow parameter that is a function of pressure and temperature corrects for the non-ideal behavior of the air. The formula for this real gas critical parameter is computed according to Johnson1.

The heated gas used for testing must be at a total pressure and stagnation temperature to mimic what the missile would experience in flight. Once the gas is heated it is again choked thereby creating an opportunity to make a redundant estimate of mass flow that can be used to identify any system leaks. The propane flow rate, prior to entering the air heater, is controlled with a set of two electro-hydraulically actuated valves mounted in parallel to provide flow rates that vary by a factor of 100. The propane flowrate is measured using a turbine flowmeter sized appropriately for the expected range of flowrates. Makeup oxygen can be added to restore the oxygen content consumed in the combustion of propane, if required, as in the case of air breathing propulsion tests. The heated gas, thus, can have either the correct mass or mole fraction of oxygen (but, not both) as found in air. The heated gas is then considered to be vitiated air. Oxygen flow is controlled using an electro-pneumatically actuated control valve. The oxygen flowrate is calculated using pressures at the entrance to the throat of another NIST traceable calibrated venturi. The operation of the air, propane and oxygen valves is digitally controlled using a PC running National Instrument�s LabView software with full proportion-integral-differential gain control loops for each system.

The digital computer system is used to provide automatic control of the airflow and total temperature. The total temperature of the air can be adjusted along with the mass flow to match hotwall heat fluxes, and thus surface temperatures, to those expected in free flight. If a free-jet nozzle is used, the control computer adjusts the air flowrate to maintain a constant pressure in the heater such that the static pressure at the exit of the free-jet nozzle is equal to the ambient pressure throughout the test. This procedure insures that the external flow field around the test item is as free as possible from shocks or expansion waves that could result from an under- or over-expanded flow. More details of matching the flight heat flux will be given later in the paper. For direct-connect engine testing the same computer control can be used to vary the total pressure and stagnation temperature to match the inflight variation in missile altitude and velocity.

B. Data Acquisition and Instrumentation Capabilities at T-Range The following is a summary of the current instrumentation and data acquisition capabilities of T-Range and the

capabilities following the ongoing upgrade process.

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Current 1) 30 General-Purpose Channels with Full Signal Conditioning (i.e. Excitation, Filtering, Amplification, Shunt / Substitution Calibration) 2) 20 Channels of Limited Signal Conditioning (Filtering and Amplification, Substitution Calibration) 3) 58 Channels of Thermocouple Conditioning, Type-K only, Linearization in Data Reduction 4) 6 Channels of Flow-Meter (Turbine type) Conditioning / Conversion 5) High-Speed Acquisition Capability: 4.8 Mbit/Second Maximum Aggregate Telemetry Data Rate, 20 Kilosample-per-Second Maximum Sustained Per-Channel Sample Rate 6) Very-High-Speed (>200 Kilosamples-per-Second) Acquisition Capability: NONE 7) 12-bit Analog-to-Digital Conversion 8) All Calibrations Done Manually Future (Operational by January 2005) 1) 120 General-Purpose Channels with Full Signal Conditioning (i.e. Excitation, Filtering, Amplification, Shunt / Substitution Calibration) 2) 40 Channels of Limited Signal Conditioning (Filtering and Amplification, Substitution Calibration) 3) 96 Channels of Thermocouple Conditioning (64 Type-K, 32 Type-C), Linearization On-The-Fly 4) 20 Channels of Flow-Meter (Turbine type) Conditioning / Conversion 5) High-Speed Acquisition Capability: 9.6 Mbit/Second Maximum Aggregate Telemetry Data Rate, 200 Kilosample-per-Second Maximum Sustained Per-Channel Sample Rate. 6) Very-High-Speed Acquisition Capability: 48 Channels of 500 Kilosample-per-Second (Sustained Per-Channel) 7) 12-bit Analog-to-Digital Conversion 8) Fully Automated System Calibration

In shunt calibration, (Item (1)), the bridge in a transducer is deliberately unbalanced by placing a set of precision resistors in parallel with the strain elements, producing a known strain output voltage. Generally, several strain steps are used so that a curve is generated that can be fit to convert the raw digital values into accurate engineering units such as pounds per square inch. Shunt calibration is widely used in instrumentation because it is quite accurate, and actually involves the transducer itself, giving a confirmation that the transducer is correctly connected and electrically functional. Substitution calibration also can be used to generate a multi-step calibration that can be curve fit, however the source is not the transducer, but rather a NIST-traceable precision voltage source that is substituted into the signal conditioning amplifier in lieu of the real transducer signal. The voltage steps used to generate the conversion from raw digital values to accurate engineering units come from the individual transducer calibration data measured by a calibration lab. The substitution technique is very accurate, but tells the test engineer nothing about the condition of the transducer at the end of the line.

To perform transducer calibrations more efficiently, T-Range is preparing to install an auto-calibration system. In the typical multi-sensor instrumentation setup, calibration can become very time consuming if one has to manually switch in the various shunt resistances and substitution voltages, which is why automated calibration systems were created. Based upon a computerized instrumentation database, with the computer cycling a switched calibration matrix through each transducer channel or group of like transducer channels while data are being collected, the auto-calibration system will be able to calibrate up to 250 transducers in a matter of minutes. Performing the same task manually would require many hours of work, and if the calibration did not go well, it would require many more hours to track down the problem and do it over. An auto-calibration system is already in use at the Center�s rocket motor static test facility, Skytop.

The auto-calibration system is also meant to address concerns about precision and accuracy. The T-Range telemetry encoders generate a 12-bit binary word for each transducer voltage "sample" taken. A 12-bit system generates a binary word that mathematically runs from zero to 4095 "counts", thus the smallest change that can be resolved is one part in 4096. This is the system precision. On a freshly calibrated system, the general-purpose channel accuracy is better than ± 1/4 percent of the full-scale transducer reading. That being said, a caveat is added; the basic transducer accuracy may be nowhere near that good. For instance, accelerometers rarely have better than +1% accuracy (and usually worse) on a known sinusoidal vibration, and thus, no matter how accurate the system is, the reduced accelerometer data will always be limited to that accuracy plus the channel�s, or + 1.25%. The way to stay accurate is through purchasing quality transducers, and performing regular system and transducer calibrations. In a manually calibrated system such as currently used at T-Range, typically the system is only recalibrated at the start of each test program, which is less than the optimum. Upon introduction of the auto-calibration system, recalibration of the system can occur on a daily basis.

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Once the analog transducer signals are generated, filtered, and amplified, they pass through a telemetry encoder (digitizer) that converts voltage levels into binary "words" representing each of those sampled levels. These binary words are appended into long, repetitive data "frames", and sent (via a fiber optic link) to our data acquisition computers and recorders in serial (sequential) format for conversion, display, and storage. The speed at which this data is transmitted is called the bit rate. We typically measure this in millions of bits-per-second. The maximum aggregate bit rate is the maximum number of bits-per-second that our telemetry encoder can generate, regardless of the number of channels. For instance, if we have an encoder that can generate 4.8 Megabits (million bits) per second, and our digitizer creates a twelve-bit "word" for each voltage sample, then we could theoretically sample 400 channels at a rate of one thousand times per second. In actuality though, we would normally sample perhaps a couple of dozen channels at 8,000 samples per second, and then another 150 or so at 1,000 samples per second. But in no case can we go beyond the 4.8 Megabit/Second maximum aggregate rate, because the encoder simply won't go any faster. This bit rate will be doubled as a result of the planned upgrades to T-Range. In addition 48 channels of very high-speed data acquisition capability will be added.

III. Description of Test Programs that use T-Range

A. Air-Breathing Engine Testing As previously stated, T-Range can provide vitiated air from 16 to 1500 psia total pressure and temperatures

from ambient to 2200 °F with hot gas flow rates up to 250 lbm/sec. These capabilities have made the facility useful for a full range of air-breathing engine testing. The air heater (sometimes called a vitiator) used for air-breathing engine testing is located on the engine thrust stand to improve thermal response time between the vitiator and test article. Thermal expansion of the air inlet pipes is minimized, which reduces the effects of temperature change in the air delivery system on engine thrust measurements. This also reduces pressure loses from the vitiator to the engine. Facilities that use remote air heaters typically run into problems transferring the heated air to the thrust stand and can reduce the accuracy in measuring thrust. T-Range uses an arrangement of inlet pipes, mounted at right angles to the line of thrust for incoming air to the vitiator that minimizes effects of this air on the thrust measurements. This arrangement is shown in Figure 1. Various load cells are available to measure static thrust. The current stand is capable of measuring thrust from 100 to 50,000 lbf. A sonic nozzle is used to calibrate the stand to verify the accuracy of the thrust measurement.

Figure 1. Arrangement of inlet pipes to minimize effect on measured thrust of incoming air to the vitiator.

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The facility was used extensively for development for the Navy�s Low Drag Ramjet and Fasthawk combustors. One use of the facility is to provide a place to test booster separation and transition from rocket-to-ramjet takeover conditions. The photograph in Figure 2 shows the rocket motor from a ducted rocket engine firing before being ejected and initiation of air-breathing engine startup during a test. With 250 lbm/sec flowrate capability, T-Range has the ability to perform freejet testing for tactical missile-sized engines. Although it is hard to see from the photograph in Figure 2, the test pictured used four freejet nozzles blowing air over the four engine inlets. Port covers in each inlet were ejected upon burnout of the solid rocket motor. The advantage of a freejet test, besides verifying port cover removal is that inlet pressure recovery margin and unstart can be verfied.

B. Aerothermal Testing The T-Range facility is routinely used to provide simulated aeroheating environments for testing materials and

missile components for systems that fly supersonically or hypersonically at or near sea level. Recent test programs have included the Evolved Sea Sparrow and the Army�s CKEM missile systems.

IV. Test Planning at T-Range and the Rest of NAWCWD Determination of the proper conditions of total temperature and pressure and hot-gas flow rates is critical to a successful test. The Aeromechanics and Thermal Analysis Branch (Code 476100D) at NAWCWD provides integrated support to assure that test conditions match expected flight profile conditions as closely as possible The analysis performed prior to aerothermal testing routinely includes a computational fluid dynamic (CFD) simulation that solves the full Navier-Stokes equations of the expected free-flight flow field comparing it to the flow field generated at T-Range. This method allows the facility to provide the highest fidelity test possible. An evaluation that includes thermal and structural analysis of a component to be tested is routinely performed to support programmatic decisions before and after a test.

An example of the use of analysis in test planning comes from a recent radome test at T-Range.

Figure 2. Freejet test of a ducted rocket motor prior to booster separation.

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The actual missile free flight speed to be simulated varied from zero to in excess of Mach 3. T-Range nozzles are

limited to approximately Mach 3. Because the largest Mach 3 nozzle at T-Range is 9.2 inches in diameter it was necessary to determine how much of the radome could be accurately tested at the facility. A CFD analysis was performed to look at the differences in surface pressure and heat flux between free flight and at T-Range. In Figure 3 geometry of the CFD problem is shown. The predicted flow field Mach number for an 8-inch base diameter radome in the exit of T-Range�s Mach 2.97 nozzle is plotted in the figure.

Comparisons of surface pressure and heat flux versus axial location are shown in Figures 4 (a) and (b). As the plots in Figure 4 demonstrate, the computer code predicted that only the first 4 inches of the radome would be heated the same as flight. Since the test engineer was only interested in the heating in the region of the nose tip the test proceeded. If the program had been interested in testing the performance of the radome at its base then T-Range may not have been used. Extensive pre-test analysis using CFD offers the project�s test engineer a preview of the expected flow field and is also useful in interpreting test data.

V. Example of How Test Conditions are Established Both CFD and engineering level software are used to help set the test conditions at T-Range.

Figure 3. Contour plot of the predicted Mach number in the nozzle flow field over a radome shape.

Nozzle Flowfield

Radome

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Using, as an example, the same test of a radome demonstrates the way analysis helps define T-Range aerothermal test conditions. The radome was approximately 16 inches long and had an 8-inch base diameter. The computational fluid dynamic (CFD) analysis described above (see Figure 3) was performed to predict the distance axially from the radome nose tip, where the T-Range test facility could be expected to match free flight heating rates. The plot in Figure 4 (b) is a comparison of heat flux on the radome between flight and T-Range at the peak heating condition. As evidenced from Figure 4 (a) and (b), flight and T-Range conditions produced a reasonable correlation over the surface with approximately a one-to-one comparison at 2 inches aft of the nose tip. Thus, this location was selected as a comparison point for both this example and the test. Considering the comparison shown in Figure 4, which is at the peak heating condition in flight, it was determined that the T-Range facility could be used to realistically simulate captive carry and free flight profiles with rapidly varying heat flux levels. To assure

Figure 4a) Comparison of surface pressure for T-Range (Nozzle) and free flight (Freestream) over an 8-inch base diameter radome.

Figure 4b) Comparison of surface heat flux for T-Range (Nozzle) and free flight (Freestream) over an 8-inch base diameter radome.

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that the surface temperatures and heat fluxes generated at T-Range closely approximated those expected from the flight trajectory, the Aeromechanics and Thermal Analysis Branch at NAWCWD provided integrated support through an iterative test design process. A key tool in measuring the aerodynamic heating conditions produced at T-Range was a steel calorimeter whose external shape approximated that of the radome to be tested. With an external surface nearly identical to the actual composite radome to be tested, the heat transfer coefficient on the calorimeter should be nearly equal to that on the test radome when the wall temperatures are equal. The calorimeter was instrumented with thermocouples mounted on both the inner and outer surfaces.

In flight, the Mach number and altitude of the missile vary with time. T-Range used a fixed geometry nozzle. Therefore, the Mach number was nearly constant throughout the run. The facility simulated the aerothermal effects of varying flight Mach number by changing the total temperature and mass flow rate of the hot gas as a function of time. Because the nozzle contour was fixed the total pressure was held constant throughout the run to ensure that the nozzle flow was properly expanded. The Aeroheating and Thermal Analysis computer code (ATAC03)2 from ITT Aerotherm was used to predict the target heating profiles at T-Range and the in-depth temperature response of the radomes in flight. The ATAC03 code uses engineering approximations to solve for both the inviscid and viscous flowfield over a surface. The results of these calculations were the spatial and time varying convective heat transfer coefficient over the radome both in flight and in the T-Range facility. The ATAC03 code also incorporated a one-dimensional (1-D) thermal solver that calculated the transient heat conduction by implicit finite difference using the boundary conditions supplied by ATAC03.

The predicted surface thermal response on the composite dome was calculated by simulating the flight trajectory in ATAC03 on a model of the composite radome. This yielded a table of time varying wall temperatures and corresponding hot wall heat fluxes. As previously stated it was decided to choose a location 2 inches axially from the radome nose tip as the location at which heating conditions at T-Range would match those in flight. A T-Range target profile was generated using ATAC03 to simulate the T-Range flow field and match the previously calculated hot wall heat fluxes on the composite radome at various points in the flight trajectory. The result of these ATAC03 computer calculations was a table of total temperatures and mass flow rates that varied with time, which T-Range attempted to match as test conditions. The T-Range facility was run with the calorimeter dome inserted in the hot gas flow field. The measured calorimeter wall temperatures were used to verify the accuracy of the ATAC03 predictions and guide the T-Range staff in setting test conditions. The actual T-Range profile of mass flow, total pressure and temperature did not exactly match the programmed values due to less than ideal valve operation. Typical plots of the commanded versus delivered airflow and propane are shown in Figures 5 and 6, respectively. The predicted calorimeter response compared well with the measured values and is shown in Figure 7. This process was repeated until the desired calorimeter response was achieved and these settings of mass flow and total temperature as a function of time became the chosen test conditions. Adjustment of the T-Range mass flow and temperature settings was repeated until the error between the delivered and desired calorimeter temperature response reached an acceptable value and then these became the test conditions.

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Figure 5. Plot of the commanded versus delivered air mass flow during an aerothermal test.

Figure 6. Plot of the commanded versus delivered propane mass flow during an aerothermal test.

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VI. Enhancements Being Performed At T-Range

A. Increased Heating Capability for Hypersonic Testing T-Range has always been the choice for testing when high heating and shear conditions are required. To

increase the T-Range capability a new air heater and nozzle have been purchased. ITT Aerotherm, a leader in the design of components for the world�s largest arc heaters, has completed

fabrication of a sudden expansion burner that uses a replaceable water-cooled liner to significantly increase the mass flow and total temperature performance of both the engine and aerothermal test legs at T-Range. The operating envelope of the new heater is shown in Figure 8. The new air heater and nozzle combination will be capable of continuous operation at 4500 °F. The nozzle will be sized to operate at Mach 3.65 and will have an exit diameter of 13.4 inches. The new heater will be capable of simulating the heating conditions of a missile flying at the Mach number and altitude shown in Figure 9.

Figure 7. Four plots showing the results of comparing measuring and predicted temperatures on a radome calorimeter for four sequential runs at T-Range.

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Figure 8. Operating envelope for the new vitiator at T-Range.

Figure 9. Increased Mach number capability of the new T-Range air heater and nozzle combination.

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Increasing the mass flow rate capability is one of the primary goals for T-Range. To handle this increased flow NAWCWD has acquired enough tanks to provide 1750 cu. ft. of additional air storage capability at the T-Range facility. This will bring the total storage capacity of air to 4650 cu. ft. All of these capabilities are expected to be available by January of 2005.

VII. Test Costs The range facilities are provided to DoD and its contractors at a reasonable cost. For example, one week of T-

Range testing costs approximately $20K and includes test hardware setup, facility matching of flight conditions based upon the expected flight profile, video and still photographic coverage, data acquisition, and data reduction. T-Range has demonstrated the ability to complete 50 or more runs in a single week. The cost per test, therefore, can be very reasonable.

VIII. Conclusions The niche that T-Range fills in industry, independent research and development and similar DoD efforts should

not be overlooked in any evaluation of the facility. The selection of T-Range in the past by a large proportion of the tactical missile community for this type of technology work on expanding the boundaries of tactical missile performance demonstrates the T-Range utility in the arena of aerothermal and engine testing.

Large numbers of IR and RF domes, for example, can be tested economically, which is critical because of the statistical nature of ceramic materials and the need to generate a meaningful database. Engine development testing that can extend over many weeks can be performed at the T-Range. The major constraint of the T-Range is in its mass flow rate that limits the size of components that can be tested, which has severely limited the number of programs that could use the range. The proposed facility enhancements of increasing the heated air flow rate from 100 to 250 lbm/s and acquiring larger supersonic nozzles are directed at overcoming this size limitation and will serve to increase T-Range�s value as a national asset.

References 1 Johnson, R.C., �Real-Gas Effects in Critical-Flow-Through Nozzles and Tabulated Thermodynamic Properties�, N65-14676, NASA Lewis Research Center, Cleveland, Ohio, January 1965. 2 �User�s Manual for the Aeroheating and Thermal Analysis Code (ATAC03),� ITT Aerotherm Final Report 0608A-02-001, January 2002.