Joseph D. Norris

Mach 10 Boundary Layer Transition Experiments on Sharp and Blunted Cones (Invited)

Boundary-layer transition and stability data were obtained at Mach 10 in the Arnold Engineering Development Complex (AEDC) Hypervelocity Wind Tunnel 9 on a 1.5-m long, 7-deg cone at unit Reynolds numbers between 1.8 and 31 million per meter. A total of 24 runs were performed at angles-of-attack between 0 and 10-deg on sharp and blunted cones with nose radii between 5.1 and 50.8-mm. The transition location was determined with coaxial thermocouples and temperature sensitive paint while stability measurements were obtained using high-frequency response pressure sensors. Mean flow and boundary layer-stability computations were also conducted and compared with the experiment. The effect of angle-of-attack and bluntness on the transition location displays similar trends compared to historical hypersonic wind tunnel data at similar Mach and Reynolds numbers. The N factor at start of transition on sharp cones increases with unit Reynolds number. Values between 4 and 7 were observed. The N factor at start of transition significantly decreases as bluntness increases and is successfully correlated with the ratio of transition location to entropy layer swallowing length. Good agreement between the computed and measured spatial amplification rates and most amplified 2 mode frequencies are obtained for sharp and moderately blunted cones. For large bluntness, where the ratio of transition to entropy swallowing length is below 0.1, 2 mode waves were not observed before the start of transition on the frustum.

Aeroheating Testing and Predictions for Project Orion CEV at Turbulent Conditions

An investigation of the aeroheating environment of the Project Orion Crew Exploration Vehicle was performed in the Arnold Engineering Development Center Hypervelocity Wind Tunnel No. 9 Mach 8 and Mach 10 nozzles and in the NASA Langley Research Center 20 - Inch Mach 6 Air Tunnel. Heating data were obtained using a thermocouple-instrumented approx.0.035-scale model (0.1778-m/7-inch diameter) of the flight vehicle. Runs were performed in the Tunnel 9 Mach 10 nozzle at free stream unit Reynolds numbers of 1x10(exp 6)/ft to 20x10(exp 6)/ft, in the Tunnel 9 Mach 8 nozzle at free stream unit Reynolds numbers of 8 x 10(exp 6)/ft to 48x10(exp 6)/ft, and in the 20-Inch Mach 6 Air Tunnel at free stream unit Reynolds numbers of 1x10(exp 6)/ft to 7x10(exp 6)/ft. In both facilities, enthalpy levels were low and the test gas (N2 in Tunnel 9 and air in the 20-Inch Mach 6) behaved as a perfect-gas. These test conditions produced laminar, transitional and turbulent data in the Tunnel 9 Mach 10 nozzle, transitional and turbulent data in the Tunnel 9 Mach 8 nozzle, and laminar and transitional data in the 20- Inch Mach 6 Air Tunnel. Laminar and turbulent predictions were generated for all wind tunnel test conditions and comparisons were performed with the experimental data to help define the accuracy of computational method. In general, it was found that both laminar data and predictions, and turbulent data and predictions, agreed to within less than the estimated 12% experimental uncertainty estimate. Laminar heating distributions from all three data sets were shown to correlate well and demonstrated Reynolds numbers independence when expressed in terms of the Stanton number based on adiabatic wall-recovery enthalpy. Transition onset locations on the leeside centerline were determined from the data and correlated in terms of boundary-layer parameters. Finally turbulent heating augmentation ratios were determined for several body-point locations and correlated in terms of the boundary-layer momentum Reynolds number.

Aerothermal Measurement Improvements using Coaxial Thermocouples at AEDC Hypervelocity Wind Tunnel No. 9

The Arnold Engineering Development Center (AEDC) Hypervelocity Wind Tunnel No. 9 Facility has played a key role in the development of hypersonic vehicles for over 30 years, providing high-quality aerodynamic and aerothermal test data covering high Mach number and high Reynolds number flight simulations. Over the last ten years, the fidelity of computational fluid dynamics tools has advanced to the point that higher accuracy validation data are needed. Therefore the experimental community must also continue to evaluate the measurement techniques utilized to provide these data, modify the techniques whenever necessary, and develop new cutting edge techniques to ensure that the highest data quality is available to support the advanced code validation efforts. An example of this process occurred recently at Tunnel 9 during two aerothermal test programs. During these tests, pretest aerothermal computations matched the general data trends reasonably well; however, an approximate 15-percent bias between the data and the computations was observed. An effort was initiated to calibrate the as-installed coaxial thermocouple gage versus a National Institute of Standards and Technology (NIST)-traceable heat-flux standard. This paper discusses the efforts taken to date and immediate recommendations to investigate the possible source of the discrepancy.

Measurements of Fluctuating Pitot Pressure, "Tunnel Noise," in the AEDC Hypervelocity Wind Tunnel No. 9

The Arnold Engineering Development Center (AEDC) Hypervelocity Wind Tunnel No. 9 facility has played a key role in the development of hypersonic vehicles for over 30 years, providing high-quality aerodynamic and aerothermal test data covering high Mach number and high Reynolds number flight simulations. Although Tunnel 9 can achieve flight level Reynolds numbers and naturally transitioning boundary layers on most test articles, the presence of “tunnel noise” can complicate the understanding of the boundary-layer transition phenomenon. In an attempt to better characterize the freestream disturbances described as “tunnel noise” a set of data was collected using a flush-mounted Pitot acoustic probe. The data quantify the relative Pitot acoustic noise of the freestream flow for the Mach 8, 10, and 14 nozzles at AEDC Tunnel 9. The percent noise level for each nozzle varied on the basis of Reynolds number from approximately 2 to 3.5 percent at Mach 8, 2.5 to 4 percent at Mach 10, and 3.75 to 6.25 percent at Mach 14.

Development of a Global Heat Transfer Measurement System at AEDC Hypervelocity Wind Tunnel 9

Heat-transfer rates are an extremely important consideration in the design of hypersonic vehicles such as atmospheric reentry vehicles. This paper describes the development of a data reduction methodology to evaluate heat-transfer rates using global surface temperature measurements on wind tunnel models at the Arnold Engineering Development Center (AEDC) White Oaks Hypervelocity Wind Tunnel 9. As a part of this development effort, a scale model of the NASA Crew Exploration Vehicle was painted with temperature-sensitive paint (TSP), and multiple sequences of high-resolution images were acquired during a five-run test program. The calculation of heat-transfer rates from TSP data acquired in Tunnel 9 is challenging because of high Reynolds number and dynamic pressure environments and the desire to use standard stainless steel wind tunnel models that were originally designed for force and moment testing. The authors developed an approach to reduce TSP data into convective heat flux while taking into consideration the challenges listed above. A preliminary comparison of the heat flux value calculated using the TSP surface temperature data with the value calculated using the standard thermocouple data is reported.

Hypersonic Global Heat-Transfer Measurements During Continuous Pitch Sweeps at AEDC Tunnel 9

The AEDC Hypervelocity Wind Tunnel No. 9 is a ground-test facility that features a unique combination of flow conditions, run times, and physical size not reproduced anywhere else in the world. By taking advantage of its large size and long run times, typical Tunnel 9 test programs are designed to collect data over a continuous angle-of-attack sweep using a wind tunnel model instrumented with a force-and-moment balance and hundreds of discrete surface pressure and heat-transfer sensors. Recently AEDC Tunnel 9 has implemented a global heat-transfer measurement system that can operate simultaneously with traditional measurement techniques during a continuous pitch sweep without significant increases in the test program’s schedule or cost. The resulting global heattransfer maps offer considerable insight into the aerothermal environment experienced by the test article. For instance, measurements of complex flow phenomena such as boundarylayer transition and shock wave boundary-layer interactions can now be mapped in detail. The data presented in this paper illustrate the advantage of global mapping for a wedge in a Mach 10 flow field at various angles of attack and Reynolds number conditions. Comparisons with data from discrete instrumentation show excellent agreement with the global measurements. The global heat-transfer measurement system developed for use in Tunnel 9 makes use of several unique innovations in illumination and camera technologies and utilizes a novel two-color temperature-sensitive paint (TSP) formulation.

The Hypervelocity Wind Tunnel No. 9 - Continued excellence through improvement and modernization

Hypervelocity Wind Tunnel No. 9, located at the White Oak, MD site of the Arnold Engineering Development Complex (AEDC), has long been recognized as a unique world class ground-test facility. The facility was developed in the early 1970s to provide critical low-altitude, high Mach number data in support of the Navys reentry development programs. Since its inception, Tunnel 9 has maintained a leading role in hypersonic ground testing by continually expanding its operational capabilities to match the needs of current and projected programs, maintaining data quality, and understanding customer requirements. Tunnel 9 started with a unique design built around a state-of-the-art supply heater that provided a clean, high-pressure, high-temperature nitrogen supply. Initial operation of Tunnel 9 realized a Mach 10 and 14 aerodynamic simulation capability. Additional Mach 7 and 8 high Reynolds number capabilities were subsequently developed. Each upgrade to Tunnel 9 during the past 40 years of operation has been in response to various sponsors or hypersonic basic research requirements. These capability enhancements have helped maintain Tunnel 9s position as a core DoD hypersonic test and evaluation (T&E) ground-test facility, which has been identified as a leading facility in all major hypersonic facility studies. Recent improvements and modernization over the past 10 years have focused on test article measurements and have significantly changed the types, quantity, and quality of test data that are readily acquired in a Tunnel 9 test entry. Recent advancements include major system changes such as a completely new control room and high-speed data system to entirely new measurement capabilities such as global heattransfer measurements using Temperature Sensitive paint technology. These along with other incremental improvements have allowed Tunnel 9 to provide new insights into the physics associated with complex hypersonic flows.

Temperature-Sensitive Paint Calibration Methodology Developed at AEDC Tunnel 9

Aerodynamic heating is a critical parameter in hypersonic vehicle design. Ground-test facilities employ a variety of discrete sensors and global techniques to characterize the heating on the surfaces of a test article. AEDC Hypervelocity Wind Tunnel No. 9 has been developing a global heat-transfer measurement technique based on temperature-sensitive paints (TSP). In order to use TSP as a heat-transfer measurement tool, a unique relationship between its emission intensity and surface temperature has to be established. A TSP calibration laboratory was developed and built at Tunnel 9. The resulting calibration process has been shown to yield repeatable calibrations insensitive to setup perturbations. The calibration laboratory has been instrumental in the development and implementation of the two-color TSP system at Tunnel 9.

Techniques for Thin-Skin Heat-Transfer Measurements for Surface Roughness at AEDC Tunnel 9 *

Arnold Engineering Development Center (AEDC) Tunnel 9 has recently developed a technique to measure transient heat-transfer rates on roughened model surfaces using the thin-skin measurement technique. The thin-skin technique is used to obtain the heat-transfer of a body by measuring the back face surface temperature using thermocouples. Tunnel 9’s short run time and high loading conditions make adapting this measurement technique challenging. A finite-element heat conduction model was created in ANSYS ® to develop the optimal thin-skin parameters for use in Tunnel 9. A novel technique for machining a wide variety of roughness elements into the wind tunnel model’s surface is discussed. This paper summarizes considerations in model design, laboratory calibration results, and recent wind tunnel data obtained using the thin-skin technique.

Use of Temperature Sensitive Paint in the AEDC Hypervelocity Wind Tunnel 9

A series of proof-of-principle tests was conducted at the Arnold Engineering Development Center (AEDC) White Oak Hypervelocity Wind Tunnel 9 to assess the feasibility of using temperature-sensitive paint (TSP) as a production test technique. A 16.75-deg, half-angle wedge model with a swept, wedge-shaped aft fin was tested at Mach 10 and Mach 14. Taking into consideration the anticipated growth rate of the instabilities introduced, approximately one-half of the model leading edge was coated with a transition grit to destabilize the boundary layer on that side of the model. Evidence of the establishment of a fully turbulent boundary layer at Mach 14 for a Reynolds number of 1.3 × 10 6 /ft is shown in the acquired TSP images. Reported data are consistent with the theory of laminar and turbulent shock wave/boundarylayer interactions with multiple embedded vortical structures. Initial qualitative results show that TSP will survive in and is a viable test technique for global surface temperature measurement in a hypervelocity blowdown facility.