Arnold Engineering Development Complex Research Articles

Numerical Investigation of Aero-Optical Distortions from High-Speed Turbulent Boundary Layers

Wall-modeled large-eddy simulations of turbulent boundary-layer flows over a flat plate at Mach 3.5, 7.87, and 13.64 were carried out, and the aero-optical distortions resulting from density fluctuations were investigated. The conditions for the Mach 3.5 case match those of a direct numerical simulation in the literature. The Mach 7.87 conditions are representative of the Hypersonic Wind Tunnel at Sandia National Laboratories, and the Mach 13.64 conditions match those of the Arnold Engineering Development Complex Hypervelocity Tunnel 9. The wall-modeled large-eddy simulations were validated using available experimental and DNS reference data. The normalized root-mean-square optical path difference for all three cases is in good agreement with respective data obtained from reference direct numerical simulations and experiments (within 1.53% for M∞=3.5, 1.25% for M∞=7.87, and 12% for M∞=13.64, compared to numerical data). Above M∞=5.0, the normalized path difference obtained from the simulations is above the value predicted by a semi-analytical relationship by Notre Dame University. With increasing Mach number, the bulk contribution to the total optical distortion shifts toward the boundary-layer edge. In addition, a proper orthogonal decomposition reveals that the nature of the dominant density fluctuations, which are located near the boundary-layer edge, changes from three-dimensional to two-dimensional. Understanding how the coherent structures contribute to the optical path difference may pave a way toward devising active flow control strategies geared toward a reduction of the optical path difference.

Evaluation of methods for characterizing the fine particulate matter emissions from aircraft and other diffusion flame combustion aerosol sources

The U. S. Environmental Protection Agency in collaboration with the U. S. Air Force Arnold Engineering Development Complex conducted the VAriable Response In Aircraft nvPM Testing (VARIAnT) 3 and 4 test campaigns to compare nonvolatile particulate matter (nvPM) emissions measurements from a variety of diffusion flame combustion aerosol sources (DFCASs), including a Cummins diesel engine, a diesel powered generator, two gas turbine start carts, a J85-GE-5 turbojet engine burning multiple fuels, and a Mini-CAST soot generator. The VARIAnT research program was devised to understand reported variability in the ARP6320A sampling system nvPM measurements. The VARIAnT research program has conducted four test campaigns to date with the VARIAnT 3 and 4 campaigns devoted to: (1) assessing the response of three different black carbon mass analyzers to particles of different size, morphology, and chemical composition; (2) characterizing the particles generated by 6 different combustion sources according to morphology, effective density, and chemical composition; and (3) assessing any significant difference between black carbon as determined by the 3 mass analyzers and the total PM determined via other techniques. Results from VARIAnT 3 and 4 campaigns revealed agreement of about 20% between the Micro-Soot Sensor, the Cavity Attenuated Phase Shift (CAPS PMSSA) monitor and the thermal-optical reference method for elemental carbon (EC) mass, independent of the calibration source used. For the LII-300, the measured mass concentrations in VARIAnT 3 fall within 18% and in VARIAnT 4 fall within 27% of the reference EC mass concentration when calibrated on a combustor rig in VARIAnT 3 and on an LGT-60 start cart in VARIAnT 4, respectively. It was also found that the three mass instrument types (MSS, CAPS PMSSA, and LII-300) can exhibit different BC to reference EC ratios depending on the emission source that appear to correlate to particle geometric mean mobility diameter, morphology, or some other parameter associated with particle geometric mean diameter (GMD) with the LII-300 showing a slightly stronger apparent trend with GMD. Systematic differences in LII-300 measured mass concentrations have been reduced by calibrating with a turbine combustion as a particle source (combustor or turbine engine). With respect to the particle size measurements, the sizing instruments (TSI SMPS, TSI EEPS, and Cambustion DMS 500) were found to be in general agreement in terms of size distributions and concentrations with some exceptions. Gravimetric measurements of the total aerosol mass produced by the various DFCAs differed from the reference EC, BC and integrated particle size distribution measured aerosol masses. The measurements of particle size distributions and single particle analysis performed using the miniSPLAT indicated the presence of larger particles (≳150 nm) having more compact morphologies, higher effective density, and a composition dominated by OC and containing ash. This increased large particle fraction is also associated with higher values of single scattering albedo measured by the CAPS PMSSA instrument and higher OC measurements. These measurements indicate gas turbine engine emissions can be a more heterogeneous mix of particle types beyond the original E−31 assumption that engine exit exhaust particles are mainly composed of black carbon.

Modal Instabilities over Blunted Cones at Angle of Attack in Hypersonic Flow

The effects of nose radius and angle of attack on the linear modal amplification over blunt circular cones at hypersonic speeds are computationally investigated. The three-dimensional laminar flow solutions over a 1.5-m-long, 7°-half-angle cone with 5.080, 9.525, and 25.40 mm nosetip radii are computed for selected angle-of-attack values and freestream conditions that match the Mach 10 experiments at a freestream unit Reynolds number of 17.1 million per meter conducted within the Hypervelocity Wind Tunnel 9 at the Air Force Arnold Engineering Development Complex (AEDC). Results indicate that the linear amplification of stationary and traveling crossflow waves along inflection lines increases with the angle of attack and decreases with the nosetip radius. The trend in transition front with respect to increasing angle of attack and bluntness is found to be consistent with the predicted increase in the amplification factors for Mack’s second mode (MM) disturbances along the streamline trajectories for the small and medium bluntness cases. The increase in the MM amplification along the windward ray for higher angles of attack is shown to be the result of a progressively earlier entropy-layer swallowing. Computations also indicate that the transition amplification factor along the windward ray increases with the angle of attack and decreases with the nosetip radius. Furthermore, the transition -factors along the leeward ray are below 2 and, therefore, rather small to lead to transition.

Supersonic Freestream Density Fluctuations from Focused Laser Differential Interferometry and Pitot-Probe Measurements

Focused laser differential interferometry (FLDI) is used to make freestream density fluctuation measurements in the von Kármán Gas Dynamics Facility (VKF) Wind Tunnel D at Arnold Engineering Development Complex. A comparison is made between pitot-probe and FLDI measurements after converting both to freestream pressure fluctuation spectra. A modification of Stainback and Wagner’s theory, incorporating recent numerical results from Chaudhry et al., is used to interpret the pitot data, while a new inversion algorithm is applied to the FLDI data. These methods of interpretation make the assumption that the tunnel freestream disturbance environment is dominated by the acoustic Kovásznay mode, in line with the theoretical and experimental evidence. Close agreement is found between the two sets of spectra, showing that accurate quantitative data can be obtained with FLDI, and the instrument can be used to extend spectra beyond the pitot bandwidth.

Assessment of a regulatory measurement system for the determination of the non-volatile particulate matter emissions from commercial aircraft engines

The SAE International has published Aerospace Information Report (AIR) 6241 which outlined the design and operation of a standardized measurement system for measuring non-volatile particulate matter (nvPM) mass and number emissions from commercial aircraft engines. Prior to this research, evaluation of this system by various investigators revealed differences in nvPM mass emissions measurement on the order of 15–30% both within a single sampling system and between two systems operating in parallel and measuring nvPM mass emissions from the same source. To investigate this issue, the U. S. Environmental Protection Agency in collaboration with the U. S. Air Force's Arnold Engineering Development Complex initiated the VAriable Response In Aircraft nvPM Testing (VARIAnT) research program to compare nvPM measurements within and between AIR-compliant sampling systems used for measuring combustion aerosols generated both by a 5201 Mini-CAST soot generator and a J85-GE-5 turbojet engine burning multiple fuels. The VARIAnT research program has conducted four test campaigns to date. The first campaign (VARIAnT 1) compared two essentially identical commercial versions of the sampling system while the second campaign (VARIAnT 2) compared a commercial system to the custom-designed Missouri University of Science and Technology's North American Reference System (NARS) built to the same specifications. Comparisons of nvPM particle mass (i.e., black carbon), number, and size were conducted in both campaigns. Additionally, the sensitivity to variation in system operational parameters was evaluated in VARIAnT 1. Results from both campaigns revealed agreement of about 12% between the two sampling systems, irrespective of manufacturer, in all aspects except for black carbon determination. The major source of measurement differences (20–70%) was due to low BC mass measurements made by the Artium Technologies LII-300 as compared to the AVL 483 Micro-Soot Sensor, the Aerodyne Cavity Attenuated Phase Shift (CAPS PMSSA) monitor, and the thermal-optical reference method for elemental carbon (EC) determination, which was used as the BC reference.

Amplification and structure of streamwise-velocity fluctuations in compression-corner shock-wave/turbulent boundary-layer interactions

In this work, we study the effect of the compression-corner angle on the streamwise turbulent kinetic energy (sTKE) and structure in Mach 2.8 flow. Krypton tagging velocimetry (KTV) is used to investigate the incoming turbulent boundary layer and flow over$8^{\circ }$,$16^{\circ }$,$24^{\circ }$and$32^{\circ }$compression corners. The experiments were performed in a 99 %$\text{N}_{2}$and 1 % Kr gas mixture in the Arnold Engineering Development Complex (AEDC) Mach 3 Calibration Tunnel (M3CT) at$Re_{\unicode[STIX]{x1D6E9}}=1750$. A figure of merit is defined as the wall-normal integrated sTKE ($\overline{\text{sTKE}}$), which is designed to identify turbulence amplification by accounting for the root-mean-squared (r.m.s.) velocity fluctuations and shear-layer width for the different geometries. We observe that the$\overline{\text{sTKE}}$increases as an exponential with the compression-corner angle near the root when normalized by the boundary-layer value. Additionally, snapshot proper orthogonal decomposition (POD) is applied to the KTV results to investigate the structure of the flow. From the POD results, we extract the dominant flow structures and compare each case by presenting mean-velocity maps that correspond to the largest positive and negative POD mode coefficients. Finally, the POD spectrum reveals an inertial range common to the boundary-layer and each compression-corner flow that is present after the first${\approx}10$dominant POD modes.

Phased acoustic array measurements of a 5.75% hybrid wing body aircraft

Detailed acoustic measurements of the noise from the leading-edge Krueger flap of a 5.75% hybrid wing body aircraft model were acquired with a traversing phased microphone array in the Arnold Engineering Development Complex NFAC 40- by 80-foot wind tunnel. The spatial resolution of the array was sufficient to distinguish between individual support brackets over the full-scale frequency range of 100 to 2875 Hz. For conditions representative of landing and take-off configuration, the noise from the brackets dominated other sources near the leading edge. Inclusion of flight-like brackets for select conditions highlights the importance of including the correct number of leading-edge high-lift device brackets with sufficient scale and fidelity. These measurements support the development of a Krueger noise model which includes cove and bracket noise.

Direct Skin Friction Measurements at Mach 10 in a Hypervelocity Wind Tunnel

This investigation concerns the design, build, and testing of direct-measuring skin friction sensors capable of performing in sustained hypersonic flow and detecting transition. A multistep approach tested the sensors through bench-test and wind-tunnel facilities. The sensors underwent National Institute of Standards and Technology traceable calibrations with well-documented uncertainties. The calibration process characterized static, thermal, pressure, and dynamic responses. Validation testing was conducted in a supersonic tunnel at Mach 4.0. The main investigations were conducted in Arnold Engineering Development Complex Tunnel 9. The skin friction sensor was integrated into a steel 155.6-cm-long (61.27 in.-long), 7 deg half-angle cone model. Flow was nominally maintained at Mach 10 and a stagnation temperature of 1250 K (2250°R). The stagnation pressure and unit Reynolds number were varied over 2.3–43.4 MPa (330–6300 psia) and (), respectively. Skin friction was measured over boundary-layer states including early transitional, transitional, and turbulent flows. Wall shear ranged between 0.92 and 340 Pa (0.02 and 7.1 psf), whereas the skin friction coefficients ranged 0.0003–0.0060. The uncertainty of the skin friction sensor remained at of the measurement for a 95% confidence level. The experimental measurements demonstrated favorable agreement with independent analyses including numerical predictions and Reynolds analogy methods.

Krypton tagging velocimetry in a turbulent Mach 2.7 boundary layer

The krypton tagging velocimetry (KTV) technique is applied to the turbulent boundary layer on the wall of the “Mach 3 Calibration Tunnel” at Arnold Engineering Development Complex (AEDC) White Oak. Profiles of velocity were measured with KTV and Pitot-pressure probes in the Mach 2.7 turbulent boundary layer comprised of 99 % $$\hbox {N}_{2}$$ /1 % Kr at momentum-thickness Reynolds numbers of $${Re}_{\varTheta }= 800, 1400$$ , and 2400. Agreement between the KTV- and Pitot-derived velocity profiles is excellent. The KTV and Pitot velocity data follow the law of the wall in the logarithmic region with application of the Van Driest I transformation. The velocity data are analyzed in the outer region of the boundary layer with the law of the wake and a velocity-defect law. KTV-derived streamwise velocity fluctuation measurements are reported and are consistent with data from the literature. To enable near-wall measurement with KTV ( $$y/\delta \approx $$ 0.1–0.2), an 800-nm longpass filter was used to block the 760.2-nm read-laser pulse. With the longpass filter, the 819.0-nm emission from the re-excited Kr can be imaged to track the displacement of the metastable tracer without imaging the reflection and scatter from the read-laser off of solid surfaces. To operate the Mach 3 AEDC Calibration Tunnel at several discrete unit Reynolds numbers, a modification was required and is described herein.

Orbital Debris Assessment Testing in the AEDC Range G

The space environment presents many hazards for satellites and spacecraft. One of the major hazards is hypervelocity impacts from uncontrolled man-made space debris. Arnold Engineering Development Complex (AEDC), the National Aeronautics and Space Administration (NASA), the United States Air Force Space and Missile Systems Center (SMC), the University of Florida, and the Aerospace Corporation configured a large ballistic range to perform a series of hypervelocity destructive impact tests in order to better understand the effects of space collisions. The test utilized AEDC's Range G light gas launcher, which is capable of firing projectiles up to 7km/sec. A nonfunctional full-scale representation of a modern satellite called the DebriSat was destroyed in the enclosed range environment. Several modifications to the range facility were made to ensure quality data was obtained from the impact events. The facility modifcations were intended to provide a high-impact energy-to-target-mass ratio (>200J/g), a nondamaging method of debris collection, and an instrumentation suite capable of providing information on the physics of the entire impact event.

Time-resolved Spectroscopy of Plasma Flash from Hypervelocity Impact on Debrisat

Time-resolved spectroscopic measurements were made on the intense plasma flash that accompanied a hypervelocity impact on “DebriSat” conducted in April 2014 at the Arnold Engineering Development Complex (AEDC) in Tennessee. The DebriSat project was a unique hypervelocity event aimed at simulating an on-orbit destructive collision of a modern satellite with a hypervelocity projectile in a single strike. Collaborators on the project included NASA's Orbital Debris Program Office at NASA Johnson Space Center (JSC), US Air Force Space and Missile Systems Center (SMC), The Aerospace Corporation, The University of Florida, and AEDC. The primary hypervelocity test object, DebriSat, was a 56kg structure representative of a modern LEO spacecraft, assembled at University of Florida following design guidance provided by Aerospace's Concept Design Center. It comprised a hexagonal body with an arrangement of seven compartmentalized bays. The main panels and the structural ribs between the bays were constructed from Al 5052 honeycomb and carbon fiber composite face sheets. Each bay had a set of structures with components intended to represent those used in a modern satellite design. The AEDC projectile was a 570g nylon-aluminum cylinder travelling at 6.8km/s. An analysis of the first several frames of the DebriSat flash shows that the rate of radial expansion of the plume decreased from 4750 m/s at t = 0 to 2250 m/s at t = 156μs. Time-resolved spectral measurements were conducted on the plasma flash with multiple frames acquired with a 10μs gating interval. Numerous emission lines were observed and were consistent with the materials of structural components located in the specific bay of DebriSat that was impacted. A background was observed in all the spectral frames and could be fitted to a blackbody temperature of ∼ 2900K. A detailed description of the plasma flash and its spectroscopic investigation is presented.

Multi-shock Shield Performance At 14 MJ for Catalogued Debris

As the orbital debris population continues to grow significant tangible threats to robotic and crewed spacecraft have risen greatly over the last decade. These threats are currently mitigated operationally; however, an understanding of engineered solutions is useful to consider with respect to the risks and costs of operational mitigation. To this end a multi-shock shield has been designed and tested to demonstrate what it takes to stop an object that fits the energy profile of catalogued debris. A 14.25MJ hypervelocity impact test has been performed on an enhanced, multi-shock shield at orbital speeds at the Arnold Engineering Development Complex (AEDC). The projectile was a hollow aluminum and nylon cylinder with characteristic dimensions typical of catalogued debris. The AEDC test of the shield occurred without any issues, and the shield successfully stopped the 598g projectile at a mass penalty of 10.35g/cm2.

Wind-Tunnel Balance Calibration with Temperature Using Design of Experiments

A statistical engineering approach featuring a formally designed experiment has recently been implemented for calibration of an internal, strain-gauged, wind-tunnel balance at Arnold Engineering Development Complex. It was necessary to consider a substantially wide variation in balance temperature for this particular application. The cost to perform a balance calibration with the balance equilibrium temperature as a completely randomized factor was therefore prohibitive. A design-of-experiments approach using a run schedule with restricted randomization, known commonly as a split-plot design, is being evaluated. For more than a decade, design of experiments has been implemented in wind-tunnel strain-gauge balance calibration processes. Calibration with design of experiments is a characterization process where data are collected and analyzed using statistical methods, allowing conclusions to be drawn with chosen levels of confidence and power. A sequential approach to new balance calibration using existing hardware was demonstrated using a crossed design. A modified, two-active-factor, Box–Behnken design was executed at three temperature levels, and a second-order regression model was constructed in all factors. When compared to the traditional approach, a minimum 50% reduction in calibration duration and a reduction in overall residual error were demonstrated.