datalama/STELLA · Datasets at Hugging Face

_id large_stringlengths 19 48	title large_stringclasses 1 value	text large_stringlengths 9 2.11k
aeronautics/07484_100.txt		e the porous cylinder is located below the model centerline,the NO flows mostly below the model centerline. A small amount of NO mixes or diffuses around to the top half ofthe flow, thereby weakly visualizing the top wake region and shear layer as well. Although not shown here, a comparison was made of two nearly identical, internally-plumbed models (Runs 4and 5) to determine the effect of varying the NO flow rate on the wake flowfield
aeronautics/14162_75.txt		m light speckles are an unpredictable variable that have a great affect on the CPC technique. The SPC technique on a larger viewing screen is superior to the CPC technique in distinguishing when the laser beam goes through the center of the simulated shock. The SPC technique results, as seen in Figure 17, are consistently better than the CPC technique. Even better results for the SPC technique can be obtained if necessary by capturing the diffraction pattern that seemed to extend beyond the camera field of view
aeronautics/21141_83.txt		xhibit, Reno, NV, Jan. 2007. 9 Broeren, A.P., Busch, G., and Bragg, M.B., “Aerodynamic Fidelity of Ice Accretion Simulation on a Subscale Model,” SAE 2007 Transactions: Journal of Aerospace, Vol. 116, Aug. 2008, pp. 560-575. 10 Broeren A.P., Diebold, J.M., and Bragg, M.B
aeronautics/23063_7.txt		t devices, such as flaps and slats, constituting the primary sources[3]. Development and maturation of viable technologies that substantially reduce the airframe component of aircraft noise is of importance to the NASA Aeronautics Research Mission Directorate (ARMD). A major effort on airframe noise reduction was initiated under the Environmentally Responsible Aviation (ERA) project. Numerous concepts for mitigating the noise produced by aircraft flaps and undercarriage were conceived and evaluated via extensive modelscale testing
aeronautics/04502_138.txt		sused, there is residual twist of about 1.2 deg over the spanof the model. No wind tunnel corrections were applied tothe data. The reference chordline used in testing theSC1094 R8 airfoil was the SC1095 chordline. Thus, theangle-of-attack data for this airfoil have been shifted by–1 deg to allow comparison with the other SC1094 R8data in this report. AIRFOIL TEST COMPARISONS Lift
aeronautics/00718_71.txt		n that in the combustor and lagged thecombustor pressure by approximately 40 degrees. This mode was also observed at the other two, lowerpower, operating conditions, although at smaller amplitudesand lower frequencies, as shown in Fig. 10. ANALYSIS OF RESULTS Comparison with Analytical Results The experimental results can be compared to theanalytical results by referencing Fig. 6, which shows thepredicted pressure spectrum, and Fig. 8, which shows theme
aeronautics/02024_128.txt		plicated chemical pathways of possible interest, and/orearly jet or late plume chemistry. The transition regimebetween plume models and 3-D global models presentsa particular challenge given the potential importance ofturbulent mixing and large species fluctuations occurring on scales below typical grid lengths, of order 100kin, which can be employed
aeronautics/20987_37.txt		n the direction of flight, Y positive to the left of the flight path, and Z positive up. With the aircraft directly overhead of the reference microphone at an altitude of 100 feet above ground level (AGL), the microphone spacing was designed to provide approximately 10° angular resolution, up to 10° below the horizon. Additional microphones provide observer angles as small as 2.4° below the horizon as shown in Table
aeronautics/23457_13.txt		(PSL) as part of a series of fundamental icecrystal icing tests in 2018. The purpose of the aerothermal calibration test was to characterize the cloudoff temperature and flow of the PSL test facility prior to performing the cloud-on icing experiments. This report presents the temperature measurements and analysis of that test day, which examined uniformity and the influence of spray bars and wall temperature effects
aeronautics/03202_67.txt		e used to express each multivariaterthogonal function in terms of ordinary multivariateolynomials The orthogonal functions are generated in aanner that allows them to be decomposed withoutbiguity into an expansion of ordinary multivariateolynomials. The orthogonalization process can bepeated using arbitrary ordinal, multivariateolynomials to generate orthogonal functions ofbitrary order in the independent variables, subjectnly to limitations related to the information containedthe data. For the FASER wind tunnel data responser
aeronautics/23075_128.txt		p. 487–526.doi:10.1146/annurev.fl.20.010188.002415.[30] El-Hady, N., “Spatial Three-Dimensional Secondary Instability of Compressible Boundary-Layer,” AIAA Journal, Vol. 29,No. 5, 1991, pp. 688–696. doi:10.2514/3.10642.[31] Ng, L., and Erlebacher,
aeronautics/10307_42.txt		m the thin film gages and phosphor. The spacing between the pressure ports was 0.20-inches from center to center. This spacing was not dense enough to guarantee that peak pressure would be measured. The number of pressure ports on the target was limited due to the small interior size of the model and the sizing of the tubes. Also, several of the thin film gages were determined to be bad and the data from these gages is not shown in the following figures. Generally
aeronautics/23913_90.txt		M with one or more auto generated, amended operation plans that divert the aircraft to alternate vertiports. The alternate vertiport play would be necessary if the VAS informs the PSU or FM of a sudden and indefinite, or extended, vertiport closure for the arrival vertiport for one or more vehicles in the FM’s fleet
aeronautics/21727_84.txt		troducing downwash acceleration is one of the key changesmade to the jet analysis to represent rotor behavior. However, reality is far different from momentum theory assumptions in that the effects of a finite number of blades,tip vortices, and root cut-out play a substantial role in theresulting non-uniform inflow distribution. Accurate representation of the inflow distribution is the first step indownwash prediction. Flow velocities measured for the isolated single rotorconfiguration are shown in Fig. 9
aeronautics/24056_268.txt		6°. The increments are smaller in the landing sizing scenario than for the takeoff L/D scenario. If larger aileron deflections are feasible, the landing scenario predicts increased AFC increment, up to 0.35% (Figure 27). An alternate approach to increasing the overall wing area to reduce approach speed is to enlarge the relative flap area. Generally, the weight implication of this approach is smaller than resizing the entire wing
aeronautics/00585_830.txt		eometries, with and without boundary-layer tripping, deployed at both moderate and high flap angles. The acoustic database isobtained from a Small Aperture Directional Array (SADA) of microphones, which was constructed to electronically steer todifferent regions of the model and to obtain farfield noise spectra and directivity from these regions. The basic flap-edgeaerodynamics is established by static surface pressure data, as well as by Computational Fluid Dynamics (CFD) calculations andsimplified
aeronautics/17969_52.txt		l.,16 the asymptotic expansion of the quantity inparentheses (for large γ) is: This expansion can be further simplified by applying the Dowell and Bliss method,15 whereby algebraicmanipulation is used to re-write each term in Eq. 27, and applying the binomial theorem. This procedureis outlined in the appendix. The resulting transform of the pressure, including terms up to O(1/M5) is:
aeronautics/03087_38.txt		F,and freestream unit Reynolds numbers from 0.01 to0.41 million per foot. A contoured axisymmetricnozzle is used to provide nominal freestream Machnumbers from 5.9 to 6.2. The nozzle exit diameter is20 inches with the flow exhausting into an open jettest section; the nozzle throat diameter is 0.466-inch.A bottom-mounted model injection system can injectmodels from a sheltered position to the tunnelcentefline in less than
aeronautics/18913_163.txt		m, C., Barad, M., Housman, J., and Kiris, C., “A Comparison of Higher-Order Shock Capturing Schemes Withinthe LAVA CFD Solver,” AIAA Paper 2014-1278, Jan 13-17, National Harbor, Maryland, 2014.24Hutcheson, F., Private communication, 2014.25Shur, M., Spalart, P., and Strelets, M., “
aeronautics/20100_26.txt		h 1,900, 1,900, and 2,450 rotary wing flight hours. The fourth participant had logged 800 hours of rotary wing flight time. Ownship error relative to the hover location was displayed on a laptop monitor, and the pilot attempted to minimize the error using a gaming joystick. The Bedford rating scale (Ref.9) was used to subjectively score each pilots spare capacity at the end of each 60-second tracking run. Dependent variables were: stick position
aeronautics/12710_250.txt		f the windmill four times.” In their 1928 summary of autogyro work (ref. 4), Glauert and Lock stated that “the only point which remains uncertain is the capacity of a gyroplane to descend vertically at a low speed, but the practical importance of this question should not be exaggerated.” Drop tests were consistent with results for airscrews, so the vertical rate of descent was expected to be double that recorded. Cierva devoted a large part of his 193
aeronautics/18818_78.txt		5 and 30 psf were also collected forsome cases. The nominal test section airspeed was Mach 0.1162. Test data consist of force, moment,and aeroelastic deflection measurements, and werecollected during α-sweep and q∞-sweep runs. Theaeroelastic deflection measurements were providedby a VICON motion tracking system. Figure 22 is aphotograph of the flexible wing wind tunnel model inthe UWAL test section during the
aeronautics/16608_93.txt		e farthest apart on the slat (S11) and the main wing (M1). Because there was only one KuliteTM transducer over the flap surface, the present measurements do not provide any information about the spanwise coherence over the flap surface. For the KuliteTM transducers mounted over the slat, the spanwise coherence had large values at the frequencies corresponding to the NBPs. This behavior was first noted using computational simulations in Ref. 11. The additional NBP
aeronautics/01788_37.txt		l speeds withreductions in drag of 20% or more shown in wind tunneltests. FIG. 5 shows a comparison of a tractor-trailer with a shortcab and a porous skin mounted directly on the trailingsurface of the trailer and extending 10% of the total lengthof the trailer forward of the trailing surface and a tractortrailer with a short cab with no porous skins.
aeronautics/23057_63.txt		s) if hr < 100 tor hs < 100 t, where t is the face tolerance. A metric is formed at the vertex with the eigenvalues of 1/h2n, 1/h2r, 1/h2sandeigenvectors of nˆ,rˆ, sˆ. The surface curvature metric is intersected with other geometry metric constraints at this vertex. An example from the Japan Aerospace Exploration Agency Standard Model (JSM) with nacelle
aeronautics/09832_88.txt		g whether the design meets requirements in many integrated vehicle areas. Many of these were described in the paper. As the Ares I design matures, with each passing month, more uses for the data output from the Monte Carlo runs are discovered as the various subsystems look for their design cases. After the design is firm, Monte Carlo simulation will be used to assist in verifying that the Ares I meets its requirements. Acknowledgments
aeronautics/04758_73.txt		H., “Numerical simulation of wake vortices measured during the Idaho Falls and Memphis field programs. 14th AIAA Applied Aerodynamics Conference, Proceedings, Part- II, 17-20 June 1996, New Orleans, LA, AIAA Paper No. 96-2496, pp. 943-960. 18Proctor, F. H., and Han, J
aeronautics/14393_71.txt		” requiring resources in excess of those available to the general population. The nascent “personal aircraft” market is HUGE, with estimates ranging up to $1T/yearWorldwide – far larger than the current transport aircraft markets in the U.S
aeronautics/23249_98.txt		e small. Now, applying BEMT yields the result that the inflow angle is calculated from where (a) is the lift curve slope of the airfoil commonly taken as 5.73 per radian and (x) is the nondimensional blade radius station, calculated as x = r/R. Knight and Hefner saw that the blade element of attack would then appear as
aeronautics/02024_48.txt		d range found in the direct numerical simulation study of Spalart and Wray [1996]. In previouswork we have found little variation of the basic wakedecay (rate of wake dispersion, overall vertical and horizontal extent, wake lifetime, etc.) with downstreamdomain size. The intcraction between a given Crow period and its neighbors is only weakly affected by thedetailed structure of those neighbors, so replacing themwith periodic images proves to be a reasonable approximation. Providing for multiple Crow
aeronautics/25132_56.txt		y discussed algorithms. With knowledge of the efficiency of each EE string, the healthier strings, i.e., the ones that require less overall power and will therefore cause less wear and tear on the GTE, can be utilized more heavily. Figure 12 shows the same scenario as Figure 11, but accounts for the efficiency of each string. One can immediately see that initially, before the failures, the EE thrust commands are separate, rather than being all the same. Also, they are not evenly spaced
aeronautics/23785_151.txt		g from the proprotor, and can ultimately lead to negative damping (i.e., instability). In order to achieve reliable predictions, extra care must be given to dynamic load cases such as proprotor whirl flutter (for a recent example with hingeless proprotors, see Ref. [33]). Inattention to seemingly minordetails, such as the alignment of structural load members, can influence stability margins
aeronautics/08866_308.txt		PCY requires that DLIC be performed to determine the aircraft’s capability to accomplish the service. The FLIPCY operational context and time sequence is shown in figures 59 and 60. Mapping of FLIPCY service to WCDMA functions is shown in figure 61. Appendix C—Summary of Previous Indepth Assessments Many of indepth technical assessments supporting the FCS technology assessment were performed during the FCS Phase II summary. An overview of the work performed and results
aeronautics/03172_25.txt		d and assessed.n objective processing method is developednd utilized to determine vortex definition. otor Model Test Description The HART II program was conducted in theen-jet, anechoic test section of the Large Loweed Facility (LLF) of the DNW, which has an exitzzle of 8 m by 6 m that provides a 19m-long freewith a low-turbulence potential core.
aeronautics/12710_503.txt		t Ames Research Center) initiated a program of theoretical and experimental research in rotor dynamics. Robert A. Ormiston came to Ames Research Center in February 1968 after conducting PhD research on the aerodynamics of a sailwing at Princeton University, where Professor Earl Dowell read his thesis. Although his original assignment at Ames was aerodynamics, all the problems and all the questions attracted him to rotor dynamics; he was in the control room of the 40- by 80-Foot Wind Tunnel
aeronautics/23077_20.txt		g Görtler instability. One of the interesting findings of the secondary instability computations by Li et al. [26] was the existence of a symmetric unstable mode that originated from the Mack mode instability of the unperturbed boundary layer and was subsequently distorted by the presence of increasingly stronger modulation of the basic flow due to the axial growth of the Görtler vortices
aeronautics/11863_62.txt		t on the 4BPF tone varied with sideline angle, with no overall noise reduction. Wake filling properties of TEB address two fan noise mechanisms. Reductions in the wake defect are expected to give reductions in rotor/stator interaction tone levels. Fan broadband noise is thought to relate to flow turbulence levels, so reductions in rotor downstream turbulence would be expected to somewhat reduce broadband noise levels. The preceding far-field acoustic analysis has shown that TEB can, with the proper flow rate, reduce both
aeronautics/08281_114.txt		lmost no loss in numericalefficiency, since we use a CFL number of 1000. Calculations were performed for Cases 1, 9, and 10 with theRK/implicit scheme using 3 stages and CFL = 1000. The solutions were obtained with the same (moderatelyhigh aspect ratio) 320 × 64 grid used for the results of the RK/implicit scheme with 5 stages. Figure 10shows the convergence histories with the matrix and
aeronautics/26388_355.txt		n aircraft design, operational efficiency, and environmental performance? AI-Driven Passenger Flow Optimization in Airports • Overview: This AI-driven system for optimizing passenger flow in airports relies on noninvasive sensors collecting biometric, psychometric, and spatio-temporal data to create a comprehensive, real-time understanding of passenger movement. The model aims to enhance operational efficiency and passenger experience through predictive congestion management and personalized guidance, while addressing privacy and ethical concerns
aeronautics/18476_99.txt		. Both traffic scenarios were developed by an ATC subject matter expert and designed to provide equivalent pilot workload. In each scenario, eight intruders were scripted to progress to a self separation then collision avoidance alert, absent of pilot action, while four different intruders were scripted to progress only to a preventive alert. All encounters were built for a single intruder, however, dynamic changes to the surrounding traffic made it possible for multiple intruders to occur simultaneously. Live traffic data was referenced in order
aeronautics/01119_78.txt		Dow ComingDC-200Fluid)graduallythinsundertheinfluenceofsurfaceshear.Undersuitableassumptions, theslopeoftheoil surfaceat theleadingedgeof theoil issimplyrelatedto thecomponentof localskinfrictionperpendicularto theedge.Opticalinterferometryprovidesasensitiveprobefor measuringthewedgeangleof thefilm.
aeronautics/08342_428.txt		Once the route is selected and modified, the ITP generates alternates for that route and a pilot task plan. If the ITP has a pilot’s log feature, it may compare the regulatory requirements for flying the mission with the capabilities and qualifications of both the aircraft and the pilot. If the pilot has flown the route before and the ITP detects changes from that previous route (e.g., different procedures at the destination airport), the ITP can alert the pilot to those differences.
aeronautics/01266_4.txt		100 with the CFD/DSMC overlap at a Knudsen nmnber of0.01 and the DSMC/collisionless DSMC overlap at 10. The present results include density contours, density and temperature along the stagnation streamline for all cases, andstagnation pressure and heat transfer coefficient as a function of Knudsen number. Nomenclature CHDKn pPRT1"X, y, ZZPHeat transfer coefficient., q'"'/_p
aeronautics/11945_61.txt		d a final state toward the end of a maximum run time period when the model has achieved thermal equilibrium. For the unconditioned state, the temperatures at the six model locations vary significantly throughout the boundary-layer survey. The other two conditioned thermal states show much less cone temperature variation at a given cone station during a boundary-layer survey and appear to be approaching a linear temperature distribution. The dashed line was designated as the thermal wall boundary condition for the CFD results finally used for comparison with the experimental measurements as
aeronautics/15300_3.txt		s. The aircraft was sized for a 6600-lb payload anda range of 300 nm. The rotor planform and twist were optimized for hover and cruise performance. For the presentrotor performance calculations, the collective pitch angle is progressively increased up to and through stall with theshaft angle set to zero. The effects of lift offset on rotor lift, power, controls, and blade airloads and structural loadsare examined
aeronautics/12917_56.txt		8, Berlin, Germany. [2] Hirt, S.M. and Anderson, B.H., “Experimental investigation of the application of microramp flow control to an oblique shock interaction”, AIAA Paper 2009-919, 47th Aerospace Sciences Meeting, January 5-8, Orlando, FL. [3] Ford, C.W.P. and Babinsky, H., “Micro-ramp control of oblique shock wave / boundary layer interactions”, AIAA
aeronautics/22224_9.txt		nfiguration. The former nomenclature, rather than the latter, will be used herein. The remote measurement approach utilizing an ITP is a reasonably well-established method for measurement underconditions that the transducers cannot directly tolerate [10, 11]. However, the experimental determination/validation ofa transfer function relating the desired unsteady pressure to the ITP-transducer measured value is necessary, particularlyif the instrumentation layout requires a long sense line or if higher frequencies (say, a couple of kHz)
aeronautics/12710_60.txt		.................................................. 33Houbolt and Brooks ..........................................
aeronautics/15387_80.txt		o 1.54e-15 and the rms of the SA model residuals varied from 2.61e-13 to 2.77e-13. Figure 24 shows the grid convergence of skin friction at a specific location on the plate and the drag coefficient integrated over the plate, both of which are computed using various USM3D options described earlier. Local skin friction and integrated drag show close agreement among various USM3D solutions on the finest two grids. It can
aeronautics/09592_73.txt		–824, 1997.7Nadarajah, S.K., Jameson, A., “Optimal Control of Unsteady Flows Using a Time Accurate Method,” AIAA Paper2002-5436, 2002.8Nadarajah, S.K., McMullen, M., and Jameson, A., “Non-Linear Frequency Domain Based Optimum Shape Design forUnsteady Three-Dimensional Flow,” AIAA Paper 2006-1052
aeronautics/09416_41.txt		d at ~420 sec). However, the LES winds result in a largest angle of attack at touchdown. Note that the parachute trims near 10 deg total angle of attack. Since the aerodynamics of the parachute dominates the capsule/parachute combination, the parachute attitude is reflected in the trim attitude of the capsule as well. Figures 17 and 18 show the pitch angle responses of the capsule and parachute respectively for each of the wind profiles. Comparison of Figs
aeronautics/20574_39.txt		tions were not producing the expected improvements in pressure distribution near the outboard side edge of the flap. The pressure history at the dynamic sensor locations was obtained for the last 34k time steps. These results were processed in the same manner as those for the enriched grid for PSD distribution over 8 overlapping blocks of 8k time steps. No surface pressure data for use with the FWH propagation code were recorded for this grid.
aeronautics/20534_38.txt		(F) sweep and a reverse (R)sweep both existed, the iterative procedure determined the optimum ∆αprb and ∆βprb such that (αprb)F + ∆αprb,(αprb)R + ∆αprb, (βprb)F + ∆βprb, and (βprb)R + ∆βprb, when rotated through ΘP to the tunnel coordinate system,would minimize the
aeronautics/19872_13.txt		s of CFD results. A variety of grid systems (both structured and unstructured) were used. Trends due to flap angle were analyzed, and effects of grid family, grid density, solver, and turbulence model were addressed. Some participants also assessed the effects of support brackets used to attach the flap and slat to the main wing for the experiment. In general, CFD results tended to under-predict lift, drag, and the magnitude of the pitching moment (moment was negative) compared with
aeronautics/23519_19.txt		bility to simulate the break-up and heat transfer effects inside the engine [22-24]. This presentstudy looks at a different engine geometry with a hidden core design. The research engine was never in production so there are not any known ice-crystal icing issues. Therefore it provided an ideal testbed for the 1DIcing Risk Analysis tool to perform an assessment to determine any potential icing risk[25]. Additionally, new icing measurement techniques were assessed to provide insight on the ice cloud state in
aeronautics/10769_87.txt		3(a). V.A.2. Spanwise Periodic Roughness Arrays: DREs The definition of a spanwise periodic array of roughness elements has at least four independent parametersthat can be optimized to accomplish a desired effect; 1) the roughness size and shape (e.g. cylindrical withradius r), 2) the roughness spacing (i.e. wavelength in spanwise direction), 3) the chordwise position onthe airfoil, and 4
aeronautics/10768_11.txt		nwell-known upwind schemes that utilize a 1D approximate Riemann solver at cell interfaces. Unique flow dependent variables determined by the Riemann solver are required at these cell interfaces to integrate the governing conservation laws over the discretized domain. When applied to multi-dimensional flows, such one-dimensional approximations give rise to undesired numerical problems, especially for triangular or tetrahedral unstructured meshes [5]
aeronautics/00428_647.txt		d tests are discussed to explain the capabilities this facility provides and to demonstrate the experience of thestaff. down Wind Tunnels; Propulsion System P erformanc e; Flight Simulation; Aerodynamic Heatin The Wall Interference Correction System (WICS) is operational at the National Transonic Facility (NTF) of NASA LangleyResearch Center (NASA LaRC) for semispan and full span tests in the solid wail (slots covered) configuration. The method isbased on
aeronautics/15680_4.txt		ssed, along with the status of FY14 efforts and future plans. I. Introduction The field of Avionics is advancing far more rapidly in terrestrial applications than in space flight applications. Spaceflight Avionics are not keeping pace with expectations set by terrestrial experience, nor are they keeping pace with the need for increasingly complex automation and crew interfaces as we move beyond Low Earth Orbit. NASA must take advantage of the strides being made by both space-related and terrestrial industries to drive our development and
aeronautics/11355_23.txt		r the downwash study, a set of three consistently sized grids with varying tail incidences weregenerated. For the tail-off configuration, the tail solid surfaces were removed, and the volume grid was re-generatedwith the same spacing requirements in the location of the tail as were used in the tail-on configuration. A summaryof the Case 1 mixed-element grid sizes and runs is shown in Table III. For the Case 3 high Reynolds number grid,the surface grid distribution from the Case 1 medium grid
aeronautics/16144_68.txt		n to have minimum pressure RMS values, to establish a base noise level (PA8410m). Five Kulite® LLE-2DC-750 pressure transducers were mounted on the TA spider arms for acoustic monitoring during the Segment 2 envelope expansion flights. A perspective view, illustrating the structure of the TA, can be seen in Fig. 10; the sensor locations are highlighted in red in Fig. 11. The pressure data from these sensors, sampled at 1000 samples per
aeronautics/19806_59.txt		tially blocked slots (Run 55) was adopted for the small separation case (case C). Another factor which can contribute to 3D effects is the fact that the channel width narrows at the upstream end of the splitter plate (from 381mm to 355mm wide) and then re-expands near the aft end of the splitter plate (from 355 mm to 381mm wide). Figure 5c shows a plan view of the splitter plate and a profile view of the
aeronautics/07420_53.txt		2 descent phase air data, aerodynamic attitude, and control deflections arepresented in appendix A. Identified on these time histories are the segments of flight corresponding to the MSdescribed in the “Mission Description” section and table 1 above. Figures A-01 through A-05 show the comparisonsof the postflight trajectory reconstruction to the INS flight data. Figure A-01 shows that the BET- and INS-basedMach numbers are generally within 1 percent of each other and that
aeronautics/12557_36.txt		xact match to the flight case, lying about halfway between the geometric scaled nozzles and flight. The effect of increasing the exit angle from 5° to 15° is shown in Fig. 6b. While not making a large difference in maximum plume diameter the initial plume region agrees much better. Lastly, note the dramatic increase in plume size as the freestream Mach number and exit pressure ratio increases (due to static pressure decreasing).6
aeronautics/08191_19.txt		TPT (to 9 x 106) Electromagnetic and piezo ZMF actuators – Actuator performance degraded at large Re due topressure – Actuators did not have sufficient authority to attach flap atrequired flap deflections • Multiple excitation locations improve performance CFD Validation of Unsteady Flows • Turbulent Separation Control of Flow over Wall-mountedHump Model (AIAA 2004-2220 Greenblatt and others)
aeronautics/15897_72.txt		....... 37Figure 56. Integral Cap Tooling Plate Installation ........................................................................ 38Figure 57
aeronautics/17943_52.txt		o data points where the engine did not experience an icing event even when the ice water content was increased. From the test data that was generated in PSL, all test data points with ice ingestion were analyzed utilizing the computer models described in Section I (Figure 3). An engine icing event caused by ice accretion was defined by the test engineers as the point where the engine thrust was observed to decrease to a value of approximately 93% of the thrust load value before the ice cloud initiation
aeronautics/22462_2.txt		yton engine with that of a Stirling engine to create a highly efficient recuperating gas turbine engine. In the explored case, both Brayton cycle and Stirling cycle engines are used to generate electrical power. Additionally, the Stirling engine is used to draw heat out of the Brayton turbine (acting to cool the turbine blades), while also pumping heat into Brayton cycle just before combustion occurs (acting as the mechanism for recuperation)
aeronautics/23830_116.txt		2006, https://doi.org/10.2514/6.2006-206.3. Grzych, M. and Mason, J., “Weather Conditions Associated with Jet Engine Power Loss and Damage due to the Ingestion of Ice Particles: What We’ve Learned Through 2009,” in AMSMeeting, Jan. 2010, https://ams.confex.com/ams/90annual/techprogram/paper_165923.htm.4. Mason, J
aeronautics/03683_54.txt		y. Accumulation Parameter, Ac The accumulation parameter directly determines the quantity of ice produced. In scaling calculations, the matching of scale and reference Ac (see eq. (4)) allows the determination of the scale spray time. In principle, time can be set easily in test facilities, so that there is no need for the scale value of this parameter to deviate from the reference. However, for short spray times, spray-system stabilization may be a significant portion of the total spray for
aeronautics/22920_337.txt		g n-alkyl chain component of the mixed C7A compositions upon ARF were temperature dependent. The C11A/C7A (50/50) and C5MEG/C7A (50/50) surfaces would be presumably in a “fluid-like” state based on the melting point of the analogous compounds (Tables 4 and 5). The 1-decanol (i.e
aeronautics/01024_289.txt		rison with two-dimensional analyses. Notethat there is no experimental data available for thisconfiguration. The parameters for this configuration are such thatthe mid-span location corresponds to a flat platecascade with a stagger angle of 45 degrees, unit gapto-chord ratio, operating in a uniform mean flow at aMach number of 0.7 parallel to the blades. The rotorhas 24 blades with a hub/tip ratio of 0.8. The radiusat the hub is
aeronautics/24261_43.txt		y. We define conflicts as when two or more flightsare projected to arrive at around the same time. For the purpose of this project, we used a window of 2 minutes. Thissituation can occur for a multitude of reasons (early or delayed flights, deviations due to weather, reduced runwaycapacity, etc.). If a conflict arises, preventative measures must be taken so that the corresponding flights arrive atdifferent times. These measures include maneuvers such as S-turns and holding patterns that elevate the
aeronautics/26345_149.txt		imented with introducing an additional angle of attack design variable for two of the drag minimization optimizations. For all cases, the lift coefficient was constrained to a specified target to ensure sufficient lift for steady level flight. A structural KS stress aggregation function was also applied to all problems, such that it would not exceed the design stress limit. Thickness bounds were set from0.1 inches to 0.3 inches
aeronautics/10979_9.txt		s chemistry, were clearly defined within the matrix of simulations performed to define the aerodatabook. Support of aerothermal design for the Orion entry capsule and more recent robotic entry vehicles to Marshave brought new simulation and validation needs to the forefront. Reaction control system jets fired fromthe base of the vehicles for active control authority interact with the high speed flow expanding around thecorner of the vehicle into the wake.[5] Simulations are executed to quantify the magnitude of jet interactioneffects on aerodynamic coefficients and
aeronautics/22970_70.txt		d SRB data has been rotated in Fig. 23(when compared to the view in Fig. 21) to get the best magnification factor for the figure. The upper surface data ofFig. 23a show symmetric flow over and between the forward port and starboard SRBs. Another area where the CFD predicts a larger absolute value of CP presence than that measured by PSP is the tip ofthe SRB nose cones as seen in the lower surface data of Fig. 21
aeronautics/22253_46.txt		xuê T\|xGê pN>(Tê 9Y+êp ê $ê xÎ«¦«Â¸ê Ypê ¹«ÎÓ«ÇÂê ÈÓê Ç£ê 2(,>ê The first calibration data set comes from a machine calibration of NASA’s MC60E balance
aeronautics/03702_52.txt		l-scale, 72-in.-chord NACA 23012 wing section. Test Facility The IWT (fig. 3) is an atmosphericpressure, closed-loop refrigerated tunnel measuring 40 ft by 70 ft overall. It has an external 200-hp electric motor driving a 79-in.-diameter axial fan to provide wind velocity, a 70-ton-capacity refrigeration system for cooling, and two 75-hp air compressors dedicated to supply air
aeronautics/25124_28.txt		g., for two quantity ofinterest vectors in the (𝛼 , 𝛽) space 2) compute the average range of residuals within data groups of interest, e.g., from a residual range space𝑅(𝛼, 𝛽) = \| ®𝑞1 − ®𝑞2\| all 𝛽 measurements may be averaged together to yield an average residual range space 𝑅¯(𝛼);3) approximate the standard deviation
aeronautics/22253_7.txt		1 = forward normal force component of a force balanceN2 = aft normal force component of a force balancePM = pitching moment of a balanceRM = rolling moment of a balanceSF = side force of a balanceS1 = forward side force component of a force balanceS2 = aft side force component of a force balanceYM = yawing moment of a balance ∆R = vector of gage output differences of an n–component balanceΘ = load iteration tolerance expressed as a percentage of
aeronautics/14019_8.txt		d of discontinuities for systems of stiff reacting flows remain a challengefor algorithm development [15]. In addition to the minimization of numerical dissipation while maintaining numericalstability in compressible turbulence with strong shocks, Yee & Sjogreen and Yee & Sweby ¨[19, 20, 16] discussed a general framework for the design of such schemes. Yee & Sjogreen ¨[22], Wang et al. [15] and references cited
aeronautics/12578_533.txt		AIAA–1984–109), 1984.48. Desplas, P.: F1 Pressurized Subsonic Wind Tunnel User’s Guide. Large Technical Facilities, Le Fauga-Mauzac Wind Tunnels Departement, ONERA, France, 1998. 49. Moëns, F.: SUNSET Project: Numerical Investigations for the Preparation of the F1 Test Campaign. ONERA Technical Report No. RT 1/12405 DA
aeronautics/10870_26.txt		e floating wind turbine system shown in FIG.1 thatincludes a structural guy post and guy cables. 5 FIG. 6 is a perspective view of another alternate embodiment of the floating wind turbine system shown in FIG.1 thatincludes a pivotable connection between the rotor blades andthe rotor hub. FIG. 7 is a perspective view of another alternate embodi ment of the floating wind turbine system shown in FIG.1 thatincludes an alternate stability arm structure.
aeronautics/18498_54.txt		3, NPRc = NPRb = 1.8, and NTRc = 3.0 are shown in Figure 6 for a range of NPRt. The data for 100 and 140 have been offset 5 and 10 dB, respectively. For NPRt = 2.1, the velocities of the two outer streams are inverted since the tertiary-stream velocity is greater than that of the bypass stream
aeronautics/13327_46.txt		, which results in bounds that quantify the potential error in the output for a given input, i.e.,a magnitude error or ‘vertical’ uncertainty. The article by Hemsch et al.13 describes in detail the processinvolved in the estimation of aerodynamic uncertainties for the Ares I launch vehicle during the ascent phaseof flight. In most instances, when the data trends only exhibit small gradients, a pure magnitude uncertaintyis largely sufficient to cover for potential phase errors, even though unaccounted
aeronautics/14703_89.txt		s, ε i (i = 0, 1, 2, 3, ..., n) , at the i-th strain stations needed for inputs to calculate the beam deflections were generated by converting the SPAR-generated axial nodal stresses, σi (i = 0, 1, 2, 3, ..., n) , at the i-th strain station into the associated bending strains, ε i , through Hooke’s law (that is, εi = σ
aeronautics/08986_60.txt		t test was simulated using LS-DYNA. The model is shown in Figure 15(a). The fuselage section model is the same as used in previous impact simulations. A 15-ft by 15-ft by 2.5-ft block of sand was modeled using solid elements. The material properties assigned to the sand were obtained from a model that was developed in 2001 for correlation with test data obtained in a drop test onto a similar type of sand (Fasanella et al.,
aeronautics/09116_9.txt		o optidly design the aconstic treatment for maximum attenuation within a specified nacelle geometry. Time domain methods are a topic of considerable research interest; however, the frequency domain method remains the primary basis for the prdction of liner attenuation in the aircraft noise industry. Although engine designs incorporate threedimensional (3D) gmmetries, a literature survey reveals that until the turn of the 2 1 st century, the only numerical method for predicting liner attenuations in
aeronautics/25822_18.txt		n the sonic boom analysis methods to compute the ground signatures of the aircraft and the Ae,r target. For any inviscid CFD solver and fixed OML with flow-through nacelles, an aircraft’s undertrack dp/p at 3BL below the aircraft is approximately invariant with respect to altitude change (see Fig. 3a) as long as the AoA [or lift coefficient (CL)] and cruise Mach number are fixed
aeronautics/19518_127.txt		m is where ∆τ =dτ ijdti ∆t is the transformed time step. To linearize Eq. (60), the following identities are used: the minimization statement in Eq. (60) by ε2 and taking the limit of ε → 0 results in The governing equation for the shadow trajectory Q0i, Eq. (61), is linearized around the reference solutionQi, Combining Eqs. (63) and (64), the following least
aeronautics/17297_42.txt		gure 9d shows this concept installed on the 18% scale aircraft model. C. 14 x 22 Wind Tunnel Acquisition of acoustic measurements and determination of the flyover directivity patterns for flap/gear noise sources required that the 14 x 22 facility be operated in an open-wall (jet) mode whereby the ceiling and the side walls of the tunnel test section are raised
aeronautics/11371_72.txt		equate to conduct V & V of complex systems and that statistical validation tools are needed. NASA Responses The responses from NASA generally discussed subjects which could provide for improved V & V but needed additional research. These subjects included: the advocacy of probability analysis as a means to translate the objectives of the system into a set of required system parameters; the need to adopt recent progress in formal methods theory and software verification tool development; and development of structural principles for mission and safety critical flight software systems.
aeronautics/12940_36.txt		e corotating primary vortices. 3.1.4 Shock-vortex interactions Although shock waves constitute a separate class of flow structures, it would be remiss not to briefly address shock-vortex interactions. At transonic speeds the presence of shocks will fundamentally alter any vortex properties also present, and one example from Schiavetta13 [2009] is shown in Figure 8
aeronautics/12671_97.txt		g the lateral aerodynamic drag characteristics. The stability derivative basis of the q" &&facilitated this analysis through the variation of terms such as to assess the impact of drag prediction on the lateral flapping in synthesized piloted Lateral Reposition maneuvers. Pilot Modeling of Lateral Reposition
aeronautics/21892_132.txt		t axis. The higher strength vortex near the groundcan be inferred from the streamtraces in Fig. 12b as wellas by comparing the Vr velocity magnitude exhibited bythe green and magenta curves shown in Fig. 21a. The single rotor with fuselage shows larger negative Vr near theground than the isolated rotor IGE. A stagnation region beneath an isolated rotor IGE has been identified in previous studies. For example, measurements on a model scale isolated single rotor (
aeronautics/10330_40.txt		. The wavelength of the narrowband dye laserthat provided the second pump beam was adjusted to 561.0 nm toposition the most prominent feature of the CO, Raman spectrum,the 00°0-10°0 band, close to the nitrogen bandhead. The abscissainFig. 3 correspondsto the Raman shiftregionfortheCO 2 molecule.InFig. 3,theRamanshiftN2 spectrum hasbeendecreasedby964cmIn fact,
aeronautics/09420_85.txt		, anisotropic grids that accommodate better alignment of stretchedtetrahedral cells with the wave angles. To demonstrate the utility of the new technology for computing sonic boom problems, a generic wing-bodyconfiguration referred to as the Segmented Leading Edge (SLE) model has been employed. The model (shown inFigure 15) has been tested in the NASA Langley Unitary Plan Supersonic Wind Tunnel (UPWT) as well as theNASA Glenn 10!10 Wind Tunnel.
aeronautics/20084_112.txt		. 23Figure 22. Simplified flowchart of airfoil adjustment code. .............................................................. 24Figure 23. Radial interpolation check of lift coefficient at M = 0.30. ................................................ 25Figure 24. Stall delay as function of angle of attack or Mach number for various radial stations
aeronautics/23313_174.txt		eraction between negotiation and flight priority. Participants argued that a flight that has priority status (e.g., due to a system failure) should not have to negotiate and should be allowed to fly through others’ space (as is the case currently in the UTM system). In effect, a priority flight would “win” every negotiation it encounters. However, this leads to two arguments
aeronautics/17994_102.txt		6. A gas turbine engine comprising the rotor blade of claim 55 1. 7. A rotor stage comprising a plurality of circumferentially arranged rotor blades as set forth in claim 1, wherein in operation of the rotor stage the exit pressure ratio profile has an absolute value of at least 1.3 for each of the relative span 6o heights from 75% to 95%. 8

aeronautics/07484_100.txt

e the porous cylinder is located below the model centerline,the NO flows mostly below the model centerline. A small amount of NO mixes or diffuses around to the top half ofthe flow, thereby weakly visualizing the top wake region and shear layer as well. Although not shown here, a comparison was made of two nearly identical, internally-plumbed models (Runs 4and 5) to determine the effect of varying the NO flow rate on the wake flowfield

aeronautics/14162_75.txt

m light speckles are an unpredictable variable that have a great affect on the CPC technique. The SPC technique on a larger viewing screen is superior to the CPC technique in distinguishing when the laser beam goes through the center of the simulated shock. The SPC technique results, as seen in Figure 17, are consistently better than the CPC technique. Even better results for the SPC technique can be obtained if necessary by capturing the diffraction pattern that seemed to extend beyond the camera field of view

aeronautics/21141_83.txt

xhibit, Reno, NV, Jan. 2007. 9 Broeren, A.P., Busch, G., and Bragg, M.B., “Aerodynamic Fidelity of Ice Accretion Simulation on a Subscale Model,” SAE 2007 Transactions: Journal of Aerospace, Vol. 116, Aug. 2008, pp. 560-575. 10 Broeren A.P., Diebold, J.M., and Bragg, M.B

aeronautics/23063_7.txt

t devices, such as flaps and slats, constituting the primary sources[3]. Development and maturation of viable technologies that substantially reduce the airframe component of aircraft noise is of importance to the NASA Aeronautics Research Mission Directorate (ARMD). A major effort on airframe noise reduction was initiated under the Environmentally Responsible Aviation (ERA) project. Numerous concepts for mitigating the noise produced by aircraft flaps and undercarriage were conceived and evaluated via extensive modelscale testing

aeronautics/04502_138.txt

sused, there is residual twist of about 1.2 deg over the spanof the model. No wind tunnel corrections were applied tothe data. The reference chordline used in testing theSC1094 R8 airfoil was the SC1095 chordline. Thus, theangle-of-attack data for this airfoil have been shifted by–1 deg to allow comparison with the other SC1094 R8data in this report. AIRFOIL TEST COMPARISONS Lift

aeronautics/00718_71.txt

n that in the combustor and lagged thecombustor pressure by approximately 40 degrees. This mode was also observed at the other two, lowerpower, operating conditions, although at smaller amplitudesand lower frequencies, as shown in Fig. 10. ANALYSIS OF RESULTS Comparison with Analytical Results The experimental results can be compared to theanalytical results by referencing Fig. 6, which shows thepredicted pressure spectrum, and Fig. 8, which shows theme

aeronautics/02024_128.txt

plicated chemical pathways of possible interest, and/orearly jet or late plume chemistry. The transition regimebetween plume models and 3-D global models presentsa particular challenge given the potential importance ofturbulent mixing and large species fluctuations occurring on scales below typical grid lengths, of order 100kin, which can be employed

aeronautics/20987_37.txt

n the direction of flight, Y positive to the left of the flight path, and Z positive up. With the aircraft directly overhead of the reference microphone at an altitude of 100 feet above ground level (AGL), the microphone spacing was designed to provide approximately 10° angular resolution, up to 10° below the horizon. Additional microphones provide observer angles as small as 2.4° below the horizon as shown in Table

aeronautics/23457_13.txt

(PSL) as part of a series of fundamental icecrystal icing tests in 2018. The purpose of the aerothermal calibration test was to characterize the cloudoff temperature and flow of the PSL test facility prior to performing the cloud-on icing experiments. This report presents the temperature measurements and analysis of that test day, which examined uniformity and the influence of spray bars and wall temperature effects

aeronautics/03202_67.txt

e used to express each multivariaterthogonal function in terms of ordinary multivariateolynomials The orthogonal functions are generated in aanner that allows them to be decomposed withoutbiguity into an expansion of ordinary multivariateolynomials. The orthogonalization process can bepeated using arbitrary ordinal, multivariateolynomials to generate orthogonal functions ofbitrary order in the independent variables, subjectnly to limitations related to the information containedthe data. For the FASER wind tunnel data responser

aeronautics/23075_128.txt

p. 487–526.doi:10.1146/annurev.fl.20.010188.002415.[30] El-Hady, N., “Spatial Three-Dimensional Secondary Instability of Compressible Boundary-Layer,” AIAA Journal, Vol. 29,No. 5, 1991, pp. 688–696. doi:10.2514/3.10642.[31] Ng, L., and Erlebacher,

aeronautics/10307_42.txt

m the thin film gages and phosphor. The spacing between the pressure ports was 0.20-inches from center to center. This spacing was not dense enough to guarantee that peak pressure would be measured. The number of pressure ports on the target was limited due to the small interior size of the model and the sizing of the tubes. Also, several of the thin film gages were determined to be bad and the data from these gages is not shown in the following figures. Generally

aeronautics/23913_90.txt

M with one or more auto generated, amended operation plans that divert the aircraft to alternate vertiports. The alternate vertiport play would be necessary if the VAS informs the PSU or FM of a sudden and indefinite, or extended, vertiport closure for the arrival vertiport for one or more vehicles in the FM’s fleet

aeronautics/21727_84.txt

troducing downwash acceleration is one of the key changesmade to the jet analysis to represent rotor behavior. However, reality is far different from momentum theory assumptions in that the effects of a finite number of blades,tip vortices, and root cut-out play a substantial role in theresulting non-uniform inflow distribution. Accurate representation of the inflow distribution is the first step indownwash prediction. Flow velocities measured for the isolated single rotorconfiguration are shown in Fig. 9

aeronautics/24056_268.txt

6°. The increments are smaller in the landing sizing scenario than for the takeoff L/D scenario. If larger aileron deflections are feasible, the landing scenario predicts increased AFC increment, up to 0.35% (Figure 27). An alternate approach to increasing the overall wing area to reduce approach speed is to enlarge the relative flap area. Generally, the weight implication of this approach is smaller than resizing the entire wing

aeronautics/00585_830.txt

eometries, with and without boundary-layer tripping, deployed at both moderate and high flap angles. The acoustic database isobtained from a Small Aperture Directional Array (SADA) of microphones, which was constructed to electronically steer todifferent regions of the model and to obtain farfield noise spectra and directivity from these regions. The basic flap-edgeaerodynamics is established by static surface pressure data, as well as by Computational Fluid Dynamics (CFD) calculations andsimplified

aeronautics/17969_52.txt

l.,16 the asymptotic expansion of the quantity inparentheses (for large γ) is: This expansion can be further simplified by applying the Dowell and Bliss method,15 whereby algebraicmanipulation is used to re-write each term in Eq. 27, and applying the binomial theorem. This procedureis outlined in the appendix. The resulting transform of the pressure, including terms up to O(1/M5) is:

aeronautics/03087_38.txt

F,and freestream unit Reynolds numbers from 0.01 to0.41 million per foot. A contoured axisymmetricnozzle is used to provide nominal freestream Machnumbers from 5.9 to 6.2. The nozzle exit diameter is20 inches with the flow exhausting into an open jettest section; the nozzle throat diameter is 0.466-inch.A bottom-mounted model injection system can injectmodels from a sheltered position to the tunnelcentefline in less than

aeronautics/18913_163.txt

m, C., Barad, M., Housman, J., and Kiris, C., “A Comparison of Higher-Order Shock Capturing Schemes Withinthe LAVA CFD Solver,” AIAA Paper 2014-1278, Jan 13-17, National Harbor, Maryland, 2014.24Hutcheson, F., Private communication, 2014.25Shur, M., Spalart, P., and Strelets, M., “

aeronautics/20100_26.txt

h 1,900, 1,900, and 2,450 rotary wing flight hours. The fourth participant had logged 800 hours of rotary wing flight time. Ownship error relative to the hover location was displayed on a laptop monitor, and the pilot attempted to minimize the error using a gaming joystick. The Bedford rating scale (Ref.9) was used to subjectively score each pilots spare capacity at the end of each 60-second tracking run. Dependent variables were: stick position

aeronautics/12710_250.txt

f the windmill four times.” In their 1928 summary of autogyro work (ref. 4), Glauert and Lock stated that “the only point which remains uncertain is the capacity of a gyroplane to descend vertically at a low speed, but the practical importance of this question should not be exaggerated.” Drop tests were consistent with results for airscrews, so the vertical rate of descent was expected to be double that recorded. Cierva devoted a large part of his 193

aeronautics/18818_78.txt

5 and 30 psf were also collected forsome cases. The nominal test section airspeed was Mach 0.1162. Test data consist of force, moment,and aeroelastic deflection measurements, and werecollected during α-sweep and q∞-sweep runs. Theaeroelastic deflection measurements were providedby a VICON motion tracking system. Figure 22 is aphotograph of the flexible wing wind tunnel model inthe UWAL test section during the

aeronautics/16608_93.txt

e farthest apart on the slat (S11) and the main wing (M1). Because there was only one KuliteTM transducer over the flap surface, the present measurements do not provide any information about the spanwise coherence over the flap surface. For the KuliteTM transducers mounted over the slat, the spanwise coherence had large values at the frequencies corresponding to the NBPs. This behavior was first noted using computational simulations in Ref. 11. The additional NBP

aeronautics/01788_37.txt

l speeds withreductions in drag of 20% or more shown in wind tunneltests. FIG. 5 shows a comparison of a tractor-trailer with a shortcab and a porous skin mounted directly on the trailingsurface of the trailer and extending 10% of the total lengthof the trailer forward of the trailing surface and a tractortrailer with a short cab with no porous skins.

aeronautics/23057_63.txt

s) if hr < 100 tor hs < 100 t, where t is the face tolerance. A metric is formed at the vertex with the eigenvalues of 1/h2n, 1/h2r, 1/h2sandeigenvectors of nˆ,rˆ, sˆ. The surface curvature metric is intersected with other geometry metric constraints at this vertex. An example from the Japan Aerospace Exploration Agency Standard Model (JSM) with nacelle

aeronautics/09832_88.txt

g whether the design meets requirements in many integrated vehicle areas. Many of these were described in the paper. As the Ares I design matures, with each passing month, more uses for the data output from the Monte Carlo runs are discovered as the various subsystems look for their design cases. After the design is firm, Monte Carlo simulation will be used to assist in verifying that the Ares I meets its requirements. Acknowledgments

aeronautics/04758_73.txt

H., “Numerical simulation of wake vortices measured during the Idaho Falls and Memphis field programs. 14th AIAA Applied Aerodynamics Conference, Proceedings, Part- II, 17-20 June 1996, New Orleans, LA, AIAA Paper No. 96-2496, pp. 943-960. 18Proctor, F. H., and Han, J

aeronautics/14393_71.txt

” requiring resources in excess of those available to the general population. The nascent “personal aircraft” market is HUGE, with estimates ranging up to $1T/yearWorldwide – far larger than the current transport aircraft markets in the U.S

aeronautics/23249_98.txt

e small. Now, applying BEMT yields the result that the inflow angle is calculated from where (a) is the lift curve slope of the airfoil commonly taken as 5.73 per radian and (x) is the nondimensional blade radius station, calculated as x = r/R. Knight and Hefner saw that the blade element of attack would then appear as

aeronautics/02024_48.txt

d range found in the direct numerical simulation study of Spalart and Wray [1996]. In previouswork we have found little variation of the basic wakedecay (rate of wake dispersion, overall vertical and horizontal extent, wake lifetime, etc.) with downstreamdomain size. The intcraction between a given Crow period and its neighbors is only weakly affected by thedetailed structure of those neighbors, so replacing themwith periodic images proves to be a reasonable approximation. Providing for multiple Crow

aeronautics/25132_56.txt

y discussed algorithms. With knowledge of the efficiency of each EE string, the healthier strings, i.e., the ones that require less overall power and will therefore cause less wear and tear on the GTE, can be utilized more heavily. Figure 12 shows the same scenario as Figure 11, but accounts for the efficiency of each string. One can immediately see that initially, before the failures, the EE thrust commands are separate, rather than being all the same. Also, they are not evenly spaced

aeronautics/23785_151.txt

g from the proprotor, and can ultimately lead to negative damping (i.e., instability). In order to achieve reliable predictions, extra care must be given to dynamic load cases such as proprotor whirl flutter (for a recent example with hingeless proprotors, see Ref. [33]). Inattention to seemingly minordetails, such as the alignment of structural load members, can influence stability margins

aeronautics/08866_308.txt

PCY requires that DLIC be performed to determine the aircraft’s capability to accomplish the service. The FLIPCY operational context and time sequence is shown in figures 59 and 60. Mapping of FLIPCY service to WCDMA functions is shown in figure 61. Appendix C—Summary of Previous Indepth Assessments Many of indepth technical assessments supporting the FCS technology assessment were performed during the FCS Phase II summary. An overview of the work performed and results

aeronautics/03172_25.txt

d and assessed.n objective processing method is developednd utilized to determine vortex definition. otor Model Test Description The HART II program was conducted in theen-jet, anechoic test section of the Large Loweed Facility (LLF) of the DNW, which has an exitzzle of 8 m by 6 m that provides a 19m-long freewith a low-turbulence potential core.

aeronautics/12710_503.txt

t Ames Research Center) initiated a program of theoretical and experimental research in rotor dynamics. Robert A. Ormiston came to Ames Research Center in February 1968 after conducting PhD research on the aerodynamics of a sailwing at Princeton University, where Professor Earl Dowell read his thesis. Although his original assignment at Ames was aerodynamics, all the problems and all the questions attracted him to rotor dynamics; he was in the control room of the 40- by 80-Foot Wind Tunnel

aeronautics/23077_20.txt

g Görtler instability. One of the interesting findings of the secondary instability computations by Li et al. [26] was the existence of a symmetric unstable mode that originated from the Mack mode instability of the unperturbed boundary layer and was subsequently distorted by the presence of increasingly stronger modulation of the basic flow due to the axial growth of the Görtler vortices

aeronautics/11863_62.txt

t on the 4BPF tone varied with sideline angle, with no overall noise reduction. Wake filling properties of TEB address two fan noise mechanisms. Reductions in the wake defect are expected to give reductions in rotor/stator interaction tone levels. Fan broadband noise is thought to relate to flow turbulence levels, so reductions in rotor downstream turbulence would be expected to somewhat reduce broadband noise levels. The preceding far-field acoustic analysis has shown that TEB can, with the proper flow rate, reduce both

aeronautics/08281_114.txt

lmost no loss in numericalefficiency, since we use a CFL number of 1000. Calculations were performed for Cases 1, 9, and 10 with theRK/implicit scheme using 3 stages and CFL = 1000. The solutions were obtained with the same (moderatelyhigh aspect ratio) 320 × 64 grid used for the results of the RK/implicit scheme with 5 stages. Figure 10shows the convergence histories with the matrix and

aeronautics/26388_355.txt

n aircraft design, operational efficiency, and environmental performance? AI-Driven Passenger Flow Optimization in Airports • Overview: This AI-driven system for optimizing passenger flow in airports relies on noninvasive sensors collecting biometric, psychometric, and spatio-temporal data to create a comprehensive, real-time understanding of passenger movement. The model aims to enhance operational efficiency and passenger experience through predictive congestion management and personalized guidance, while addressing privacy and ethical concerns

aeronautics/18476_99.txt

. Both traffic scenarios were developed by an ATC subject matter expert and designed to provide equivalent pilot workload. In each scenario, eight intruders were scripted to progress to a self separation then collision avoidance alert, absent of pilot action, while four different intruders were scripted to progress only to a preventive alert. All encounters were built for a single intruder, however, dynamic changes to the surrounding traffic made it possible for multiple intruders to occur simultaneously. Live traffic data was referenced in order

aeronautics/01119_78.txt

Dow ComingDC-200Fluid)graduallythinsundertheinfluenceofsurfaceshear.Undersuitableassumptions, theslopeoftheoil surfaceat theleadingedgeof theoil issimplyrelatedto thecomponentof localskinfrictionperpendicularto theedge.Opticalinterferometryprovidesasensitiveprobefor measuringthewedgeangleof thefilm.

aeronautics/08342_428.txt

Once the route is selected and modified, the ITP generates alternates for that route and a pilot task plan. If the ITP has a pilot’s log feature, it may compare the regulatory requirements for flying the mission with the capabilities and qualifications of both the aircraft and the pilot. If the pilot has flown the route before and the ITP detects changes from that previous route (e.g., different procedures at the destination airport), the ITP can alert the pilot to those differences.

aeronautics/01266_4.txt

100 with the CFD/DSMC overlap at a Knudsen nmnber of0.01 and the DSMC/collisionless DSMC overlap at 10. The present results include density contours, density and temperature along the stagnation streamline for all cases, andstagnation pressure and heat transfer coefficient as a function of Knudsen number. Nomenclature CHDKn pPRT1"X, y, ZZPHeat transfer coefficient., q'"'/_p

aeronautics/11945_61.txt

d a final state toward the end of a maximum run time period when the model has achieved thermal equilibrium. For the unconditioned state, the temperatures at the six model locations vary significantly throughout the boundary-layer survey. The other two conditioned thermal states show much less cone temperature variation at a given cone station during a boundary-layer survey and appear to be approaching a linear temperature distribution. The dashed line was designated as the thermal wall boundary condition for the CFD results finally used for comparison with the experimental measurements as

aeronautics/15300_3.txt

s. The aircraft was sized for a 6600-lb payload anda range of 300 nm. The rotor planform and twist were optimized for hover and cruise performance. For the presentrotor performance calculations, the collective pitch angle is progressively increased up to and through stall with theshaft angle set to zero. The effects of lift offset on rotor lift, power, controls, and blade airloads and structural loadsare examined

aeronautics/12917_56.txt

8, Berlin, Germany. [2] Hirt, S.M. and Anderson, B.H., “Experimental investigation of the application of microramp flow control to an oblique shock interaction”, AIAA Paper 2009-919, 47th Aerospace Sciences Meeting, January 5-8, Orlando, FL. [3] Ford, C.W.P. and Babinsky, H., “Micro-ramp control of oblique shock wave / boundary layer interactions”, AIAA

aeronautics/22224_9.txt

nfiguration. The former nomenclature, rather than the latter, will be used herein. The remote measurement approach utilizing an ITP is a reasonably well-established method for measurement underconditions that the transducers cannot directly tolerate [10, 11]. However, the experimental determination/validation ofa transfer function relating the desired unsteady pressure to the ITP-transducer measured value is necessary, particularlyif the instrumentation layout requires a long sense line or if higher frequencies (say, a couple of kHz)

aeronautics/12710_60.txt

.................................................. 33Houbolt and Brooks ..........................................

aeronautics/15387_80.txt

o 1.54e-15 and the rms of the SA model residuals varied from 2.61e-13 to 2.77e-13. Figure 24 shows the grid convergence of skin friction at a specific location on the plate and the drag coefficient integrated over the plate, both of which are computed using various USM3D options described earlier. Local skin friction and integrated drag show close agreement among various USM3D solutions on the finest two grids. It can

aeronautics/09592_73.txt

–824, 1997.7Nadarajah, S.K., Jameson, A., “Optimal Control of Unsteady Flows Using a Time Accurate Method,” AIAA Paper2002-5436, 2002.8Nadarajah, S.K., McMullen, M., and Jameson, A., “Non-Linear Frequency Domain Based Optimum Shape Design forUnsteady Three-Dimensional Flow,” AIAA Paper 2006-1052

aeronautics/09416_41.txt

d at ~420 sec). However, the LES winds result in a largest angle of attack at touchdown. Note that the parachute trims near 10 deg total angle of attack. Since the aerodynamics of the parachute dominates the capsule/parachute combination, the parachute attitude is reflected in the trim attitude of the capsule as well. Figures 17 and 18 show the pitch angle responses of the capsule and parachute respectively for each of the wind profiles. Comparison of Figs

aeronautics/20574_39.txt

tions were not producing the expected improvements in pressure distribution near the outboard side edge of the flap. The pressure history at the dynamic sensor locations was obtained for the last 34k time steps. These results were processed in the same manner as those for the enriched grid for PSD distribution over 8 overlapping blocks of 8k time steps. No surface pressure data for use with the FWH propagation code were recorded for this grid.

aeronautics/20534_38.txt

(F) sweep and a reverse (R)sweep both existed, the iterative procedure determined the optimum ∆αprb and ∆βprb such that (αprb)F + ∆αprb,(αprb)R + ∆αprb, (βprb)F + ∆βprb, and (βprb)R + ∆βprb, when rotated through ΘP to the tunnel coordinate system,would minimize the

aeronautics/19872_13.txt

s of CFD results. A variety of grid systems (both structured and unstructured) were used. Trends due to flap angle were analyzed, and effects of grid family, grid density, solver, and turbulence model were addressed. Some participants also assessed the effects of support brackets used to attach the flap and slat to the main wing for the experiment. In general, CFD results tended to under-predict lift, drag, and the magnitude of the pitching moment (moment was negative) compared with

aeronautics/23519_19.txt

bility to simulate the break-up and heat transfer effects inside the engine [22-24]. This presentstudy looks at a different engine geometry with a hidden core design. The research engine was never in production so there are not any known ice-crystal icing issues. Therefore it provided an ideal testbed for the 1DIcing Risk Analysis tool to perform an assessment to determine any potential icing risk[25]. Additionally, new icing measurement techniques were assessed to provide insight on the ice cloud state in

aeronautics/10769_87.txt

3(a). V.A.2. Spanwise Periodic Roughness Arrays: DREs The definition of a spanwise periodic array of roughness elements has at least four independent parametersthat can be optimized to accomplish a desired effect; 1) the roughness size and shape (e.g. cylindrical withradius r), 2) the roughness spacing (i.e. wavelength in spanwise direction), 3) the chordwise position onthe airfoil, and 4

aeronautics/10768_11.txt

nwell-known upwind schemes that utilize a 1D approximate Riemann solver at cell interfaces. Unique flow dependent variables determined by the Riemann solver are required at these cell interfaces to integrate the governing conservation laws over the discretized domain. When applied to multi-dimensional flows, such one-dimensional approximations give rise to undesired numerical problems, especially for triangular or tetrahedral unstructured meshes [5]

aeronautics/00428_647.txt

d tests are discussed to explain the capabilities this facility provides and to demonstrate the experience of thestaff. down Wind Tunnels; Propulsion System P erformanc e; Flight Simulation; Aerodynamic Heatin The Wall Interference Correction System (WICS) is operational at the National Transonic Facility (NTF) of NASA LangleyResearch Center (NASA LaRC) for semispan and full span tests in the solid wail (slots covered) configuration. The method isbased on

aeronautics/15680_4.txt

ssed, along with the status of FY14 efforts and future plans. I. Introduction The field of Avionics is advancing far more rapidly in terrestrial applications than in space flight applications. Spaceflight Avionics are not keeping pace with expectations set by terrestrial experience, nor are they keeping pace with the need for increasingly complex automation and crew interfaces as we move beyond Low Earth Orbit. NASA must take advantage of the strides being made by both space-related and terrestrial industries to drive our development and

aeronautics/11355_23.txt

r the downwash study, a set of three consistently sized grids with varying tail incidences weregenerated. For the tail-off configuration, the tail solid surfaces were removed, and the volume grid was re-generatedwith the same spacing requirements in the location of the tail as were used in the tail-on configuration. A summaryof the Case 1 mixed-element grid sizes and runs is shown in Table III. For the Case 3 high Reynolds number grid,the surface grid distribution from the Case 1 medium grid

aeronautics/16144_68.txt

n to have minimum pressure RMS values, to establish a base noise level (PA8410m). Five Kulite® LLE-2DC-750 pressure transducers were mounted on the TA spider arms for acoustic monitoring during the Segment 2 envelope expansion flights. A perspective view, illustrating the structure of the TA, can be seen in Fig. 10; the sensor locations are highlighted in red in Fig. 11. The pressure data from these sensors, sampled at 1000 samples per

aeronautics/19806_59.txt

tially blocked slots (Run 55) was adopted for the small separation case (case C). Another factor which can contribute to 3D effects is the fact that the channel width narrows at the upstream end of the splitter plate (from 381mm to 355mm wide) and then re-expands near the aft end of the splitter plate (from 355 mm to 381mm wide). Figure 5c shows a plan view of the splitter plate and a profile view of the

aeronautics/07420_53.txt

2 descent phase air data, aerodynamic attitude, and control deflections arepresented in appendix A. Identified on these time histories are the segments of flight corresponding to the MSdescribed in the “Mission Description” section and table 1 above. Figures A-01 through A-05 show the comparisonsof the postflight trajectory reconstruction to the INS flight data. Figure A-01 shows that the BET- and INS-basedMach numbers are generally within 1 percent of each other and that

aeronautics/12557_36.txt

xact match to the flight case, lying about halfway between the geometric scaled nozzles and flight. The effect of increasing the exit angle from 5° to 15° is shown in Fig. 6b. While not making a large difference in maximum plume diameter the initial plume region agrees much better. Lastly, note the dramatic increase in plume size as the freestream Mach number and exit pressure ratio increases (due to static pressure decreasing).6

aeronautics/08191_19.txt

TPT (to 9 x 106) Electromagnetic and piezo ZMF actuators – Actuator performance degraded at large Re due topressure – Actuators did not have sufficient authority to attach flap atrequired flap deflections • Multiple excitation locations improve performance CFD Validation of Unsteady Flows • Turbulent Separation Control of Flow over Wall-mountedHump Model (AIAA 2004-2220 Greenblatt and others)

aeronautics/15897_72.txt

....... 37Figure 56. Integral Cap Tooling Plate Installation ........................................................................ 38Figure 57

aeronautics/17943_52.txt

o data points where the engine did not experience an icing event even when the ice water content was increased. From the test data that was generated in PSL, all test data points with ice ingestion were analyzed utilizing the computer models described in Section I (Figure 3). An engine icing event caused by ice accretion was defined by the test engineers as the point where the engine thrust was observed to decrease to a value of approximately 93% of the thrust load value before the ice cloud initiation

aeronautics/22462_2.txt

yton engine with that of a Stirling engine to create a highly efficient recuperating gas turbine engine. In the explored case, both Brayton cycle and Stirling cycle engines are used to generate electrical power. Additionally, the Stirling engine is used to draw heat out of the Brayton turbine (acting to cool the turbine blades), while also pumping heat into Brayton cycle just before combustion occurs (acting as the mechanism for recuperation)

aeronautics/23830_116.txt

2006, https://doi.org/10.2514/6.2006-206.3. Grzych, M. and Mason, J., “Weather Conditions Associated with Jet Engine Power Loss and Damage due to the Ingestion of Ice Particles: What We’ve Learned Through 2009,” in AMSMeeting, Jan. 2010, https://ams.confex.com/ams/90annual/techprogram/paper_165923.htm.4. Mason, J

aeronautics/03683_54.txt

y. Accumulation Parameter, Ac The accumulation parameter directly determines the quantity of ice produced. In scaling calculations, the matching of scale and reference Ac (see eq. (4)) allows the determination of the scale spray time. In principle, time can be set easily in test facilities, so that there is no need for the scale value of this parameter to deviate from the reference. However, for short spray times, spray-system stabilization may be a significant portion of the total spray for

aeronautics/22920_337.txt

g n-alkyl chain component of the mixed C7A compositions upon ARF were temperature dependent. The C11A/C7A (50/50) and C5MEG/C7A (50/50) surfaces would be presumably in a “fluid-like” state based on the melting point of the analogous compounds (Tables 4 and 5). The 1-decanol (i.e

aeronautics/01024_289.txt

rison with two-dimensional analyses. Notethat there is no experimental data available for thisconfiguration. The parameters for this configuration are such thatthe mid-span location corresponds to a flat platecascade with a stagger angle of 45 degrees, unit gapto-chord ratio, operating in a uniform mean flow at aMach number of 0.7 parallel to the blades. The rotorhas 24 blades with a hub/tip ratio of 0.8. The radiusat the hub is

aeronautics/24261_43.txt

y. We define conflicts as when two or more flightsare projected to arrive at around the same time. For the purpose of this project, we used a window of 2 minutes. Thissituation can occur for a multitude of reasons (early or delayed flights, deviations due to weather, reduced runwaycapacity, etc.). If a conflict arises, preventative measures must be taken so that the corresponding flights arrive atdifferent times. These measures include maneuvers such as S-turns and holding patterns that elevate the

aeronautics/26345_149.txt

imented with introducing an additional angle of attack design variable for two of the drag minimization optimizations. For all cases, the lift coefficient was constrained to a specified target to ensure sufficient lift for steady level flight. A structural KS stress aggregation function was also applied to all problems, such that it would not exceed the design stress limit. Thickness bounds were set from0.1 inches to 0.3 inches

aeronautics/10979_9.txt

s chemistry, were clearly defined within the matrix of simulations performed to define the aerodatabook. Support of aerothermal design for the Orion entry capsule and more recent robotic entry vehicles to Marshave brought new simulation and validation needs to the forefront. Reaction control system jets fired fromthe base of the vehicles for active control authority interact with the high speed flow expanding around thecorner of the vehicle into the wake.[5] Simulations are executed to quantify the magnitude of jet interactioneffects on aerodynamic coefficients and

aeronautics/22970_70.txt

d SRB data has been rotated in Fig. 23(when compared to the view in Fig. 21) to get the best magnification factor for the figure. The upper surface data ofFig. 23a show symmetric flow over and between the forward port and starboard SRBs. Another area where the CFD predicts a larger absolute value of CP presence than that measured by PSP is the tip ofthe SRB nose cones as seen in the lower surface data of Fig. 21

aeronautics/22253_46.txt

xuê T|xGê pN>(Tê 9Y+êp ê $ê xÎ«¦«Â¸ê Ypê ¹«ÎÓ«ÇÂê ÈÓê Ç£ê 2(,>ê The first calibration data set comes from a machine calibration of NASA’s MC60E balance

aeronautics/03702_52.txt

l-scale, 72-in.-chord NACA 23012 wing section. Test Facility The IWT (fig. 3) is an atmosphericpressure, closed-loop refrigerated tunnel measuring 40 ft by 70 ft overall. It has an external 200-hp electric motor driving a 79-in.-diameter axial fan to provide wind velocity, a 70-ton-capacity refrigeration system for cooling, and two 75-hp air compressors dedicated to supply air

aeronautics/25124_28.txt

g., for two quantity ofinterest vectors in the (𝛼 , 𝛽) space 2) compute the average range of residuals within data groups of interest, e.g., from a residual range space𝑅(𝛼, 𝛽) = | ®𝑞1 − ®𝑞2| all 𝛽 measurements may be averaged together to yield an average residual range space 𝑅¯(𝛼);3) approximate the standard deviation

aeronautics/22253_7.txt

1 = forward normal force component of a force balanceN2 = aft normal force component of a force balancePM = pitching moment of a balanceRM = rolling moment of a balanceSF = side force of a balanceS1 = forward side force component of a force balanceS2 = aft side force component of a force balanceYM = yawing moment of a balance ∆R = vector of gage output differences of an n–component balanceΘ = load iteration tolerance expressed as a percentage of

aeronautics/14019_8.txt

d of discontinuities for systems of stiff reacting flows remain a challengefor algorithm development [15]. In addition to the minimization of numerical dissipation while maintaining numericalstability in compressible turbulence with strong shocks, Yee & Sjogreen and Yee & Sweby ¨[19, 20, 16] discussed a general framework for the design of such schemes. Yee & Sjogreen ¨[22], Wang et al. [15] and references cited

aeronautics/12578_533.txt

AIAA–1984–109), 1984.48. Desplas, P.: F1 Pressurized Subsonic Wind Tunnel User’s Guide. Large Technical Facilities, Le Fauga-Mauzac Wind Tunnels Departement, ONERA, France, 1998. 49. Moëns, F.: SUNSET Project: Numerical Investigations for the Preparation of the F1 Test Campaign. ONERA Technical Report No. RT 1/12405 DA

aeronautics/10870_26.txt

e floating wind turbine system shown in FIG.1 thatincludes a structural guy post and guy cables. 5 FIG. 6 is a perspective view of another alternate embodiment of the floating wind turbine system shown in FIG.1 thatincludes a pivotable connection between the rotor blades andthe rotor hub. FIG. 7 is a perspective view of another alternate embodi ment of the floating wind turbine system shown in FIG.1 thatincludes an alternate stability arm structure.

aeronautics/18498_54.txt

3, NPRc = NPRb = 1.8, and NTRc = 3.0 are shown in Figure 6 for a range of NPRt. The data for 100 and 140 have been offset 5 and 10 dB, respectively. For NPRt = 2.1, the velocities of the two outer streams are inverted since the tertiary-stream velocity is greater than that of the bypass stream

aeronautics/13327_46.txt

, which results in bounds that quantify the potential error in the output for a given input, i.e.,a magnitude error or ‘vertical’ uncertainty. The article by Hemsch et al.13 describes in detail the processinvolved in the estimation of aerodynamic uncertainties for the Ares I launch vehicle during the ascent phaseof flight. In most instances, when the data trends only exhibit small gradients, a pure magnitude uncertaintyis largely sufficient to cover for potential phase errors, even though unaccounted

aeronautics/14703_89.txt

s, ε i (i = 0, 1, 2, 3, ..., n) , at the i-th strain stations needed for inputs to calculate the beam deflections were generated by converting the SPAR-generated axial nodal stresses, σi (i = 0, 1, 2, 3, ..., n) , at the i-th strain station into the associated bending strains, ε i , through Hooke’s law (that is, εi = σ

aeronautics/08986_60.txt

t test was simulated using LS-DYNA. The model is shown in Figure 15(a). The fuselage section model is the same as used in previous impact simulations. A 15-ft by 15-ft by 2.5-ft block of sand was modeled using solid elements. The material properties assigned to the sand were obtained from a model that was developed in 2001 for correlation with test data obtained in a drop test onto a similar type of sand (Fasanella et al.,

aeronautics/09116_9.txt

o optidly design the aconstic treatment for maximum attenuation within a specified nacelle geometry. Time domain methods are a topic of considerable research interest; however, the frequency domain method remains the primary basis for the prdction of liner attenuation in the aircraft noise industry. Although engine designs incorporate threedimensional (3D) gmmetries, a literature survey reveals that until the turn of the 2 1 st century, the only numerical method for predicting liner attenuations in

aeronautics/25822_18.txt

n the sonic boom analysis methods to compute the ground signatures of the aircraft and the Ae,r target. For any inviscid CFD solver and fixed OML with flow-through nacelles, an aircraft’s undertrack dp/p at 3BL below the aircraft is approximately invariant with respect to altitude change (see Fig. 3a) as long as the AoA [or lift coefficient (CL)] and cruise Mach number are fixed

aeronautics/19518_127.txt

m is where ∆τ =dτ ijdti ∆t is the transformed time step. To linearize Eq. (60), the following identities are used: the minimization statement in Eq. (60) by ε2 and taking the limit of ε → 0 results in The governing equation for the shadow trajectory Q0i, Eq. (61), is linearized around the reference solutionQi, Combining Eqs. (63) and (64), the following least

aeronautics/17297_42.txt

gure 9d shows this concept installed on the 18% scale aircraft model. C. 14 x 22 Wind Tunnel Acquisition of acoustic measurements and determination of the flyover directivity patterns for flap/gear noise sources required that the 14 x 22 facility be operated in an open-wall (jet) mode whereby the ceiling and the side walls of the tunnel test section are raised

aeronautics/11371_72.txt

equate to conduct V & V of complex systems and that statistical validation tools are needed. NASA Responses The responses from NASA generally discussed subjects which could provide for improved V & V but needed additional research. These subjects included: the advocacy of probability analysis as a means to translate the objectives of the system into a set of required system parameters; the need to adopt recent progress in formal methods theory and software verification tool development; and development of structural principles for mission and safety critical flight software systems.

aeronautics/12940_36.txt

e corotating primary vortices. 3.1.4 Shock-vortex interactions Although shock waves constitute a separate class of flow structures, it would be remiss not to briefly address shock-vortex interactions. At transonic speeds the presence of shocks will fundamentally alter any vortex properties also present, and one example from Schiavetta13 [2009] is shown in Figure 8

aeronautics/12671_97.txt

g the lateral aerodynamic drag characteristics. The stability derivative basis of the q" &&facilitated this analysis through the variation of terms such as to assess the impact of drag prediction on the lateral flapping in synthesized piloted Lateral Reposition maneuvers. Pilot Modeling of Lateral Reposition

aeronautics/21892_132.txt

t axis. The higher strength vortex near the groundcan be inferred from the streamtraces in Fig. 12b as wellas by comparing the Vr velocity magnitude exhibited bythe green and magenta curves shown in Fig. 21a. The single rotor with fuselage shows larger negative Vr near theground than the isolated rotor IGE. A stagnation region beneath an isolated rotor IGE has been identified in previous studies. For example, measurements on a model scale isolated single rotor (

aeronautics/10330_40.txt

. The wavelength of the narrowband dye laserthat provided the second pump beam was adjusted to 561.0 nm toposition the most prominent feature of the CO, Raman spectrum,the 00°0-10°0 band, close to the nitrogen bandhead. The abscissainFig. 3 correspondsto the Raman shiftregionfortheCO 2 molecule.InFig. 3,theRamanshiftN2 spectrum hasbeendecreasedby964cmIn fact,

aeronautics/09420_85.txt

, anisotropic grids that accommodate better alignment of stretchedtetrahedral cells with the wave angles. To demonstrate the utility of the new technology for computing sonic boom problems, a generic wing-bodyconfiguration referred to as the Segmented Leading Edge (SLE) model has been employed. The model (shown inFigure 15) has been tested in the NASA Langley Unitary Plan Supersonic Wind Tunnel (UPWT) as well as theNASA Glenn 10!10 Wind Tunnel.

aeronautics/20084_112.txt

. 23Figure 22. Simplified flowchart of airfoil adjustment code. .............................................................. 24Figure 23. Radial interpolation check of lift coefficient at M = 0.30. ................................................ 25Figure 24. Stall delay as function of angle of attack or Mach number for various radial stations

aeronautics/23313_174.txt

eraction between negotiation and flight priority. Participants argued that a flight that has priority status (e.g., due to a system failure) should not have to negotiate and should be allowed to fly through others’ space (as is the case currently in the UTM system). In effect, a priority flight would “win” every negotiation it encounters. However, this leads to two arguments

aeronautics/17994_102.txt

6. A gas turbine engine comprising the rotor blade of claim 55 1. 7. A rotor stage comprising a plurality of circumferentially arranged rotor blades as set forth in claim 1, wherein in operation of the rotor stage the exit pressure ratio profile has an absolute value of at least 1.3 for each of the relative span 6o heights from 75% to 95%. 8