SCIENTIFIC AND TECHNICAL AEROSPACE REPORTS
A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 44, ISSUE 14 - JULY 18, 2006
02 AERODYNAMICSIncludes aerodynamics of flight vehicles, test bodies, airframe components and combinations, wings, and control surfaces.
Also includes aerodynamics of rotors, stators, fans, and other elements of turbomachinery.
For related information see also 34 Fluid Mechanics and Thermodynamics.
20060019534 Moscow M. V. Lomonosov State Univ., Moscow, Russian Federation
Plasma Aerodynamic Experiments
Aleksandrov, A F; Bychkov, V L; Timofeev, I B; Jun 22, 2004; 25 pp.; In English; Original contains color illustrations Report No.(s): AD-A446478; No Copyright; ONLINE: http://hdl.handle.net/100.2/ADA446478; Avail.: CASI: A03, Hardcopy
As is well known plasma technologies nowadays find wide applications in aerodynamics, and it has become evident that development of hypersonic aviation is unlikely possible without application of plasma and beam technologies.
In external aerodynamics can be applied for improvement of combustion processes due to usage of created volumetric plasma structures in the stagnation zone and directly in subsonic and supersonic air-fuel mixtures. Electrical gas discharges are promising methods of combustion improvement because of their ability to produce concentrated zones with high content of chemical active species.
Actually this puts the problem of understanding the nontrivial interaction of hydrodynamic, chemical and plasma-chemical phenomena. So there is absolute necessity for detailed physical study, both experimental and theoretical, ofdischarge activated reacting flows at well-defined conditions. DTIC
Aerodynamics; Combustion; Gas Discharges; Plasmas (Physics)
20060020151 Government Accountability Office, Washington, DC, USA
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Defense Acquisitions: Better Acquisition Strategy Needed for Successful Development of the Army's Warrior Unmanned Aircraft System
May 2006; 32 pp.; In English Report No.(s): PB2006-112266; GAO-06-593; No Copyright; Avail.: CASI: A03, Hardcopy
Through 2011, the Department of Defense (DOD) plans to spend $20 billion on unmanned aircraft systems, including the Army's 'Warrior.'
Because of congressional concerns that some systems have been more costly and taken more time to produce than predicted, GAO reviewed the Warrior program.
This report (1) describes the Army's requirements underlying its decision to acquire Warrior instead of existing systems such as the Air Force's Predator, and (2) assesses whether the Army has established a sound acquisition strategy for the Warrior program.
NTIS
Air Defense; Pilotless Aircraft
20060020181 NASA Langley Research Center, Hampton, VA, USA
Planar laser-induced fluorescence (PLIF) investigation of hypersonic flowfields in a Mach 10 wind tunnel
Danehy, Paul M.; Wilkes, Jennifer A.; Aderfer, David W.; Jones, Stephen B.; Robbins, Anthony W.; Pantry, Danny P.; Schwartz, Richard J.; 2006; 15 pp.; In English; 25th AIAA Aerodynamic Measurement Technology and Ground Testing Conference, 5-8 Jun. 2006, San Francisco, CA, USA; Original contains color illustrations Contract(s)/Grant(s): WBS 489-02-07-07 Report No.(s): AIAA Pasper 2006-3442; No Copyright; Avail.: CASI: A03, Hardcopy
Planar laser-induced fluorescence (PLIF) of nitric oxide (NO) was used to visualize four different hypersonic flowfields in the NASA Langley Research Center 31-Inch Mach 10 Air wind tunnel. The four configurations were: (1) the wake flowfield of a fuselage-only X-33 lifting body, (2) flow over a flat plate containing a rectangular cavity, (3) flow over a 70deg bluntedcone with a cylindrical afterbody, formerly studied by an AGARD working group, and (4) an Apollo-geometry entry capsule- relevant to the Crew Exploration Vehicle currently being developed by NASA. In all cases, NO was seeded into the flowfieldthrough tubes inside or attached to the model sting and strut. PLIF was used to visualize the NO in the flowfield. In some cases pure NO was seeded into the flow while in other cases a 5% NO, 95% N2 mix was injected. Several parameters were varied including seeding method and location, seeding mass flow rate, model angle of attack and tunnel stagnation pressure, which varies the unit Reynolds number. The location of the laser sheet was as also varied to provide three dimensional flow information. Virtual Diagnostics Interface (ViDI) technology developed at NASA Langley was used to visualize the data sets in post processing. The measurements demonstrate some of the capabilities of the PLIF method for studying hypersonic flows. Author
Laser Induced Fluorescence; Hypersonic Flow; Nitric Oxide; Flat Plates; Flow Distribution; Afterbodies; Fuselages; Angle of Attack
20060020224 NASA Langley Research Center, Hampton, VA, USA
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Fundamental Mixing and Combustion Experiments for Propelled Hypersonic Flight, Chaper 7
Diskin, G. S.; Danehy, P. M.; Drummond, J. P.; Cutler, A. D.; [2002]; 15 pp.; In English; Original contains color and black and white illustrations Contract(s)/Grant(s): NCC1-370; 708-72-80-01; No Copyright; Avail.: CASI: A03, Hardcopy
The first experiment is a study of a coaxial jet discharging into stagnant laboratory air, with center jet of a mixture of 5% oxygen and 95% helium by volume and coflow jet of air. The exit flow pressure of both center-jet and coflow nozzles is atmosphere. The presence of oxygen in the center jet is to allow the use of an oxygen flow-tagging technique (RELIEF4) to obtain non-intrusive velocity measurements. Both jets are nominally Mach 1.8, but, because of the greater speed of sound, the center jet velocity is more than twice that of the coflow. The mixing layer which forms between the center jet and the coflow near the nozzle exit is compressible, with a calculated convective Mach number of approximately 0.7.
This geometry has several advantages: The streamwise development of the flow is generally dominated by turbulent stresses (rather than pressure forces), and thus calculations are sensitive to turbulence modeling. It includes features present in supersonic combustors, including a compressible mixing layer near the nozzle exit and a light-gas/air plume downstream. Since it is a free jet, it provides easy access for both optical instrumentation and probes. Since it is axisymmetric, it requires fewer experimental measurements to fully characterize, and calculations can be performed with more modest computer resources. However, weak shock waves formed at the nozzle exit strengthen and turn normal as they approach the axis, complicating the flow. Care is thus taken in the design of the facility to provide as near as possible to 1-D flow at the exit of both center and coflow nozzles, and to minimize the strength of waves generated at the nozzle exit. Results from this experiment are compared to CFD solutions obtained by VULCAN, a previously developed code used in engine analysis. The second experiment is a study of a supersonic combustor consisting of a diverging duct with single downstream-angled wall injector.
Thus, the geometry is relatively simple and large regions of subsonic recirculating flow are avoided. The nominal entrance Mach number is 2 and the enthalpy of the test gas (hot air 'simulant') is nominally that of Mach 7 flight. It was believed, on the basis of calculations performed that this would produce mixing-limited flow, that is to say, one for which chemical reaction to equilibrium proceeds at a much greater rate than mixing. It later proved that this was not the case. The primary experimental technique employed is coherent anti-Stokes Raman spectroscopy, known by its acronym CARS. The species probed is molecular nitrogen and the quantity measured is temperature. Intrusive probes, such as Pitot, total temperature, hot-wire, etc., are not used due to access difficulty and high heat flux in the combustor, and because they may alter the flow. CARS has several advantages over other optical methods. It is a relatively mature and well-understood technique. Signal levels are relatively high and the signal is in the form of a coherent (laser) beam that can be collected through small windows. Incoherent (non-CARS) interferences are rejected by spatial filtering. Derived from text
Computational Fluid Dynamics; Hypersonic Flight; Turbulence Models; Mixing Layers (Fluids); Supersonic Combustion Ramjet Engines
20060020688 NASA Langley Research Center, Hampton, VA, USA
Cascaded Perforates as One-Dimensional, Bulk Absorbers
Parrott, T. L.; Jones, M. G.; [2006]; 20 pp.; In English; 12th AIAA/CEAS Aeroacoustics Conference, 8-10 May 2006, Cambridge, MA, USA; Original contains color illustrations Contract(s)/Grant(s): 23-781-30-14 Report No.(s): AIAA Paper 2006-2402; No Copyright; Avail.: CASI: A03, Hardcopy
Porous cell honeycomb liners for aircraft engine nacelles offer the possibility of exploiting extended reaction effects to improve liner attenuation bandwidth as generally attributed to the performance of bulk absorbers. This paper describes an analytical procedure, starting with an impedance prediction model for a single perforated plate, to estimate the bulk-absorber parameters for a cascade of such perforates - a first step to modeling a porous wall honeycomb structure. The objective is to build confidence in a lumped element impedance model, when applied to a uniformly-spaced set of porous plates to predict its .bulk. absorber properties. The model is based upon a modified version of the two-parameter flow resistance model of the form A + BV(sub inc), where A and B are physics-based, semi-empirical parameters that are adjusted to provide an optimum fit to a composite dataset from three plate porosities of 2.5, 5 and 10%. The composite dataset is achieved by reformulating the two-parameter flow resistance model into a .reduced pressure drop coefficient. dependency on perforate hole Reynolds number. The resulting impedance model is employed to calculate surface impedance spectra for N and 2N-layer perforate cascades. The well-known two-thickness method for experimental determination of bulk-absorber parameters is then applied to these .synthesized. data sets to predict the characteristic impedance and propagation constant for the perforate cascades. These results are then compared with experimental results reported in a companion paper. Author
Honeycomb Structures; Porous Plates; Perforated Plates; Plates (Structural Members); Surface Properties; Nacelles; Flow Resistance; Reynolds Number; Linings
20060020707 NASA Langley Research Center, Hampton, VA, USA
Aerothermodynamic Analyses of Towed Ballutes 
Gnoffo, Peter A.; Buck, Greg; Moss, James N.; Nielsen, Eric; Berger, Karen; Jones, William T.; Rudavsky, Rena; [2006]; 18 pp.; In English; 9th AIAA/ASME Joint Thermophysics and Heat Transfer Conference, 5-8 Jun. 2006, San Francisco, CA, USA; Original contains color and black and white illustrations Contract(s)/Grant(s): NAS8-02130; 759-07-05 Report No.(s): AIAA Paper 2006-3771; Copyright; Avail.: CASI: A03, Hardcopy
A ballute (balloon-parachute) is an inflatable, aerodynamic drag device for application to planetary entry vehicles. Two challenging aspects of aerothermal simulation of towed ballutes are considered. The first challenge, simulation of a complete system including inflatable tethers and a trailing toroidal ballute, is addressed using the unstructured-grid, Navier-Stokes solver FUN3D. Auxiliary simulations of a semi-infinite cylinder using the rarefied flow, Direct Simulation Monte Carlo solver, DSV2, provide additional insight into limiting behavior of the aerothermal environment around tethers directly exposed to the free stream. Simulations reveal pressures higher than stagnation and corresponding large heating rates on the tether as it emerges from the spacecraft base flow and passes through the spacecraft bow shock. The footprint of the tether shock on the toroidal ballute is also subject to heating amplification. Design options to accommodate or reduce these environments are discussed. The second challenge addresses time-accurate simulation to detect the onset of unsteady flow interactions as a function of geometry and Reynolds number. Video of unsteady interactions measured in the Langley Aerothermodynamic Laboratory 20-Inch Mach 6 Air Tunnel and CFD simulations using the structured grid, Navier-Stokes solver LAURA are compared for flow over a rigid spacecraft-sting-toroid system. The experimental data provides qualitative information on the amplitude and onset of unsteady motion which is captured in the numerical simulations. The presence of severe unsteady fluid - structure interactions is undesirable and numerical simulation must be able to predict the onset of such motion. Author
Aerothermodynamics; Ballutes; Simulation; Tethering; Flow Characteristics; Design Analysis; Parachutes; Inflatable Structures
20060020753 NASA Langley Research Center, Hampton, VA, USA
Characteristic Lifelength of Coherent Structure in the Turbulent Boundary Layer
Palumbo, Daniel L.; [2006]; 16 pp.; In English; 12th AIAA/CEAS Aeroacoustics Conference, 8-10 May 2006, Cambridge, MA, USA; Original contains color illustrations Contract(s)/Grant(s): 23-781-10-13 Report No.(s): AIAA Paper 2006-2410; No Copyright; Avail.: CASI: A03, Hardcopy
A characteristic lifelength is defined by which a Gaussian distribution is fit to data correlated over a 3 sensor array sampling streamwise sidewall pressure. The data were acquired at subsonic, transonic and supersonic speeds aboard a Tu-144. Lifelengths are estimated using the cross spectrum and are shown to compare favorably with Efimtsov's prediction of correlation space scales. Lifelength distributions are computed in the time/frequency domain using an interval orrelation technique on the continuous wavelet transform of the original time data. The median values of the lifelength distributions are found to be very close to the frequency averaged result. The interval correlation technique is shown to allow the retrieval and inspection of the original time data of each event in the lifelength distribution, thus providing a means to locate and study the nature of the coherent structure in the turbulent boundary layer. The lifelength data can be converted to lifetimes using the convection velocity. The lifetime of events in the time/frequency domain are displayed in Lifetime Maps. The primary purpose of the paper is to validate these new analysis techniques so that they can be used with confidence to further characterize coherent structure in the turbulent boundary layer. Author
Tu-144 Aircraft; Turbulent Boundary Layer; Correlation; Normal Density Functions; Wavelet Analysis; Subsonic Speed; Supersonic Speed; Transonic Speed
20060020820 Air Force Inst. of Tech., Wright-Patterson AFB, OH USA
Characterizing and Controlling the Effects of Differential Drag on Satellite Formations
Wedekind, James T; Mar 2006; 127 pp.; In English Report No.(s): AD-A446933; AFIT/GSS/ENY/06-M14; No Copyright; ONLINE: http://hdl.handle.net/100.2/ADA446933; Avail.: CASI: A07, Hardcopy The ability to fly satellites in close formations represents a capability that could revolutionize the way satellite missionsare designed in the future. This study examines three of the primary formation flying designs and characterizes the effects that an anomalous satellite with a slightly different cross-sectional area would have on the stability of the formation. Following the characterization of these effects, a controller is implemented to mitigate the cross-sectional area differences between the satellites. The results show that, with the addition of a straightforward controller, small cross-sectional area differences can be mitigated and corrected such that the satellites will remain in close proximity and, in some cases, the formation will remain stable. DTIC
Aerodynamic Drag; Drag; Flight Control; Formation Flying; Satellite Constellations; Stability
20060020862 Test Group (0046th), Holloman AFB, NM USA
Holloman High Speed Test Track Design Manual
Mar 1, 2005; 83 pp.; In English; Original contains color illustrations Report No.(s): AD-A446998; AAC/PA-07-13-05-270; XC-46TW; No Copyright; ONLINE: http://hdl.handle.net/100.2/ADA446998; Avail.: CASI: A05, Hardcopy
The design of rocket sleds requires the engineer to evaluate complex loads and numerous load conditions which are imposed on the sleds as they travel along the track. This manual represents the combined effort of many Test Track engineers. In essence they have tried to document the current sled design practices that have lead to successful sled tests. This manual along with more current technical investigations and reports serve as a guide for designing sleds and sled tests. While this design manual provides adequate design guidance for most typical efforts, it can not cover all the possible scenarios. Because many of the design practices are based on experience rather than purely a scientific approach, the Flight Chief and senior Track Management take the prerogative to approve deviation from the guidelines stated in this design manual on a case-by-case basis based on their engineering knowledge and vast sled test experience. DTIC
High Speed; Manuals; Sleds
Source: NASA
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