SEARCH · Search NASA

Results for “TAIL MOUNTING”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

211 records · Page 12

All sources NASA NTRS DOE OSTI

NASA NTRS1975

Results of a space shuttle vehicle ferry configuration afterbody fairing optimization study using a 140A/B 0.0405-scale model orbiter (43-0) in the Rockwell International 7.75 by 11.0 ft low speed wind tunnel (OA124)

Experimental aerodynamic investigations were conducted on a dual-strut mounted 0.0405-scale representation of the 140A/B outer mold line space shuttle orbiter vehicle. The tests, conducted from 11 Oct., 1974 through 22 Oct., 1974, were primarily to investigate aerodynamic stability and control characteristics of the space shuttle orbiter ferry configuration. Four afterbody fairing configurations and various additions to them in the form of horizontal and ventral fins strakes and other aerodynamic protuberances were tested. Base line data on the basic orbiter with MPS nozzles and bodyflap were recorded. The drag of the optimum ferry configuration was increased to the level of the basic orbiter for possible flight test configurations by the addition of two sizes of perforated speed brakes on the tail cone surface.

Houlihan, S. R.↗

NASA NTRS1987

Plasma parameters in the near wake of the Space Shuttle

Results are presented from a Spherical Retarding Potential Analyzer and a spherical Langmuir Probe which were flown on the Spacelab-2 mission of the Space Shuttle Orbiter in August 1985. The probes were mounted in such a position that for certain Orbiter orientations, the wake edge formed by the starboard payload bay sill swept across the probes at a roll rate of once per orbit about an inertially stabilized pointing direction. Profiles of O(+) number density and electron number density across the wake edge showed that the charge particle number densities declined into the wake at a faster rate than that predicted by theoretical models. At angles of more than 10 deg into the wake from its geometric edge, two electron populations are found, with different mean random energies and electrical potentials. One of these populations is suggested to be the result of selective filtering of the high energy tails of the ambient ionospheric population, the other, a result of photoemission from the payload bay surfaces.

Raitt, W. J.↗

NASA NTRS1993

L-Band Mobile Terminal Antennas for Helicopters

The feasibility of using a low gain antenna (LGA) as a mobile terminal antenna for helicopters is described in this paper. The objectives are 1) to select the lowest cost antenna system which can be easily mounted on a helicopter and capable of communicating with a satellite, and 2) to determine the best antenna position on the helicopter to mitigate the signal blockage due to rotor blades and the multipath effect from the helicopter's body. The omni-directional low gain antenna (LGA) is selected because it is simple, reliable, and low cost. The helix antenna is selected among the many LGAs because it is the most economical one and has the widest elevation beamwidth. Both 2-arm and 4-arm helices are studied experimentally to determine the antenna's performance and the scattering effects from the helicopter's body. It is found that the LGA should be located near the tail section and at least 8 inches above the helicopter.

helicopter↗

NASA NTRS2016

Crash Testing and Simulation of a Cessna 172 Aircraft: Hard Landing Onto Concrete

A full-scale crash test of a Cessna 172 aircraft was conducted at the Landing and Impact Research Facility at NASA Langley Research Center during the summer of 2015. The purpose of the test was to evaluate the performance of Emergency Locator Transmitters (ELTs) that were mounted at various locations in the aircraft and to generate impact test data for model validation. A finite element model of the aircraft was developed for execution in LSDYNA to simulate the test. Measured impact conditions were 722.4-in/s forward velocity and 276-in/s vertical velocity with a 1.5deg pitch (nose up) attitude. These conditions were intended to represent a survivable hard landing. The impact surface was concrete. During the test, the nose gear tire impacted the concrete, followed closely by impact of the main gear tires. The main landing gear spread outward, as the nose gear stroked vertically. The only fuselage contact with the impact surface was a slight impact of the rearmost portion of the lower tail. Thus, capturing the behavior of the nose and main landing gear was essential to accurately predict the response. This paper describes the model development and presents test-analysis comparisons in three categories: inertial properties, time sequence of events, and acceleration and velocity time-histories.

Jackson, Karen E.↗

NASA NTRS1994

Wind Tunnel Testing for the Stratospheric Observatory for Infrared Astronomy

NASA Ames Research Center is pursuing the development of SOFIA, the Stratospheric Observatory For Infrared Astronomy. SOFIA will consist of a 2.5 meter telescope mounted aft of the wing of a Boeing 747 aircraft. Since a large portion of the infrared spectrum is not visible at ground level due to absorption by water vapor in the atmosphere below 40,000 feet, it is highly desirable to make observations above this altitude. SOFIA will provide the opportunity for astronomers to conduct high-altitude research for extended periods of time. Current study is focused on wind tunnel testing for the open cavity. If not controlled, air would create resonance and damage the telescope. For this reason, SOFIA will design a boundary layer control device to achieve laminar flow over the cavity. This also provides a clearer flow for seeing, thus improving resolution on infrared sources. Other effects being tested in the wind tunnel are aerodynamic torque loads on the telescope, and flutter loads on the tail.

Schenberger, Deborah↗

NASA NTRSDate not supplied

Simulations of a Turbulent Flow Subjected to Favorable and Adverse Pressure Gradients

This paper reports the results from a direct numerical simulation of an initially turbulent boundary layer passing over a wall-mounted “speed bump” geometry. The speed bump, represented in the form of a Gaussian distribution profile, generates a favorable pressure gradient region over the upstream half of the geometry, followed by an adverse pressure gradient over the downstream half. The boundary layer approaching the bump undergoes strong acceleration in the favorable pressure gradient region before experiencing incipient or very weak separation within the adverse pressure gradient region. These types of flows have proven to be particularly challenging to predict using lower-fidelity simulation tools based on various turbulence modeling approaches and warrant the use of the highest-fidelity simulation techniques. Simulation results are utilized to examine the key phenomena present in the flowfield, such as relaminarization/stabilization in the strong acceleration region succeeded by retransition to turbulence near the onset of adverse pressure gradient, incipient/weak separation, and development of internal layers where the sense of streamwise pressure gradient changes at the foot, apex and tail of the bump. The present direct numerical simulation is performed using a flow solver developed exclusively for graphics processing units, which is found to provide a significant speedup compared to an earlier solver optimized for central processing unit architectures.

Ali Uzun↗

NASA NTRS1959

Dynamic Longitudinal and Directional Stability Derivatives for a 45 deg. Sweptback-Wing Airplane Model at Transonic Speeds

An investigation was made at transonic speeds to determine some of the dynamic stability derivatives of a 45 deg. sweptback-wing airplane model. The model was sting mounted and was rigidly forced to perform a single-degree-of-freedom angular oscillation in pitch or yaw of +/- 2 deg. The investigation was made for angles of attack alpha, from -4 deg. to 14 deg. throughout most of the transonic speed range for values of reduced-frequency parameter from 0.015 to 0.040 based on wing mean aerodynamic chord and from 0.04 to 0.14 based on wing span. The results show that reduced frequency had only a small effect on the damping-in-pitch derivative and the oscillatory longitudinal stability derivative for all Mach numbers M and angles of attack with the exception of the values of damping coefficient near M = 1.03 and alpha = 8 deg. to 14 deg. In this region, the damping coefficient changed rapidly with reduced frequency and negative values of damping coefficient were measured at low values of reduced frequency. This abrupt variation of pitch damping with reduced frequency was a characteristic of the complete model or wing-body-vertical-tail combination. The damping-in-pitch derivative varied considerably with alpha and M for the horizontal-tail-on and horizontal-tail-off configurations, and the damping was relatively high at angles of attack corresponding to the onset of pitch-up for both configurations. The damping-in-yaw derivative was generally independent of reduced frequency and M at alpha = -4 deg. to 4 deg. At alpha = 8 deg. to 14 deg., the damping derivative increased with an increase in reduced frequency and alpha for the configurations having the wing, whereas the damping derivative was either independent of or decreased with increase in reduced frequency for the configuration without the wing. The oscillatory directional stability derivative for all configurations generally decreased with an increase in the reduced-frequency parameter, and, in some instances, unstable values were measured for the model configuration with the horizontal tail removed.

Bielat, Ralph P.↗

NASA NTRS2015

System Noise Assessment of Blended-Wing-Body Aircraft With Open Rotor Propulsion

An aircraft system noise study is presented for the Blended-Wing-Body (BWB) aircraft concept with three open rotor engines mounted on the upper surface of the airframe. It is shown that for such an aircraft, the cumulative Effective Perceived Noise Level (EPNL) is about 24 dB below the current aircraft noise regulations of Stage 4. While this makes the design acoustically viable in meeting the regulatory requirements, even with the consideration of more stringent noise regulations of a possible Stage 5 in the next decade or so, the design will likely meet stiff competitions from aircraft with turbofan engines. It is shown that the noise levels of the BWB design are held up by the inherently high noise levels of the open rotor engines and the limitation on the shielding benefit due to the practical design constraint on the engine location. Furthermore, it is shown that the BWB design has high levels of noise from the main landing gear, due to their exposure to high speed flow at the junction between the center body and outer wing. These are also the reasons why this baseline BWB design does not meet the NASA N+2 noise goal of 42 dB below Stage 4. To identify approaches that may further reduce noise, parametric studies are also presented, including variations in engine location, vertical tail and elevon variations, and airframe surface acoustic liner treatment effect. These have the potential to further reduce noise but they are only at the conceptual stage.

Guo, Yueping↗

DOE OSTI2020

AEROICESTUDY-Colorado State University Ice Spectrometer

The dataset comprises measures, using Colorado State University's Ice Spectrometer (IS, an immersion freezing device with a range from 0 °C down to -26 to -29 °C) of atmospheric ice nucleating particle (INP) concentrations taken at ARM's Southern Great Plains (SGP) Site Guest Instrument Facility (GIF), in Lamont, Oklahoma, during the Aerosol-Ice Formation Closure Pilot Study (AEROICESTUDY). A filter sampler was mounted approximately 6 m above ground level atop the GIF rooftop platform. Single-use filter units open to the atmosphere fitted with pre-cleaned and pre-sterilized, 47-mm diameter Nuclepore polycarbonate filters were used. Filters were typically drawn for either a full day period (9AM-5PM CDT) or broken up into morning (9AM-12PM CDT) and afternoon (1PM-5PM CDT) periods. Additionally, 4 overnight filters were taken (6PM-8AM CDT). The volume of air filtered was typically between 3000 and 8000 standard liters for daily samples. After collection, filters were stored at -20 °C until processed. Initial processing to obtain spectra of INP number concentration active via the immersion freezing mechanism versus temperature was conducted using CSU's IS instrument (McCluskey et al., 2018). For measurement of INPs, collected aerosol particles were re-suspended in 8 mL of 0.1 µm-filtered deionized (DI) water. Aliquots of each suspension, and serial dilutions, were dispensed into trays which were fit into aluminum blocks in the IS. Samples were cooled at 0.33 °C per min and the freezing temperatures of wells recorded automatically. Cumulative INP concentrations were determined by first calculating the INPs per mL of suspension based on Vali (1971) and then converting to concentration per standard liter of air using the proportion of the total liquid sample dispensed and the air sample volumes. Two-tailed, 95% confidence intervals for binomial sampling were based on Equation 2 from Agresti and Coull (1998). Aliquots of suspensions from selected samples were also heat treated (95 °C for 20 min) to denature and deactiv

54 ENVIRONMENTAL SCIENCES↗

NASA NTRS1995

Helmet-Mounted Display Research Capabilities of the NASA/Army Rotorcraft Aircrew Systems Concepts Airborne Laboratory (RASCAL)

The Rotorcraft Aircrew Systems Concepts Airborne Laboratory (RASCAL) is a UH-60A Black Hawk helicopter that is being modified by the US Army and NASA for flight systems research. The principal systems that are being installed in the aircraft are a Helmet Mounted Display (HMD) and imaging system, and a programmable full authority Research Flight Control System (RFCS). In addition, comprehensive instrumentation of both the rigid body of the helicopter and the rotor system is provided. The paper will describe the capabilities of these systems and their current state of development. A brief description of initial research applications is included. The wide (40 X 60 degree) field-of-view HMD system has been provided by Kaiser Electronics. It can be configured as a monochromatic system for use in bright daylight conditions, a two color system for darker ambients, or a full color system for use in night viewing conditions. Color imagery is achieved using field sequential video and a mechanical color wheel. In addition to the color symbology, high resolution computer-gene rated imagery from an onboard Silicon Graphics Reality Engine Onyx processor is available for research in virtual reality applications. This synthetic imagery can also be merged with real world video from a variety of imaging systems that can be installed easily on the front of the helicopter. These sensors include infrared or tv cameras, or potentially small millimeter wave radars. The Research Flight Control System is being developed for the aircraft by a team of contractors led by Boeing Helicopters. It consists of a full authority high bandwidth fly-by-wire actuators that drive the main rotor swashplate actuators and the tail rotor actuator in parallel. This arrangement allows the basic mechanical flight control system of the Black Hawk to be retained so that the safety pilot can monitor the operation of the system through the action of his own controls. The evaluation pilot will signal the fly-by-wire actuators through the flight computer from electrical sidearm controllers located in the right hand cockpit. The system will have very substantial input/output capacity and impressive computational power. These systems are installed in the aircraft using predominantly a MIL-STD 1553B data bus architecture. Sensor data from the RFCS, the basic aircraft and rotor system instrumentation including navigation information, and the HMD system are easily exchanged among user systems, or are available at the systems operator station located in the cabin for real time monitoring or data recording.

Jacobsen, R. A.↗

DOE OSTI2024

MOSAiC-Colorado State University Ice Spectrometer

This data set contains atmospheric ice nucleating particle (INP) measurements, using Colorado State University’s (CSU) Ice Spectrometer (IS), of filter collections taken at the U.S. DOE ARM AMF2 site onboard the R/V Polarstern P-deck during the Multidisciplinary Drifting Observatory for the Study of Arctic Climate (MOSAiC) field campaign. Samples were collected from October 27, 2019 to September 24, 2020. A filter sampler was mounted approximately 15 m above ground level on a railing in proximity to (and approximately 3 m below) the Aerosol Observation System (AOS) inlet. Single-use filter units open to the atmosphere were pre-cleaned and pre-loaded with 47-mm diameter Nuclepore polycarbonate (0.2 µm pore-diameter) filters. Filters were typically drawn for a three-day period, with an average volume of air filtered of 87,000 standard liters. Total volumes were calculated through recorded daily flow rates using a mass flow meter (TSI). After collection, filters were stored and transported frozen until analysis using CSU’s IS instrument (McCluskey et al., 2018). Aerosol particles were first re-suspended in 8 mL of 0.1 µm-filtered deionized (DI) water. Aliquots of each suspension, and corresponding 11-fold dilutions, were dispensed into polymerase chain reaction (PCR) trays and placed into the aluminum blocks of the IS. Samples were cooled at approximately 0.33 °C min -1 and freezing detected optically with corresponding temperatures recorded. Cumulative INP concentrations were determined through calculating the number of INPs per mL of suspension (Vali, 1971) and converting to concentration per standard L of air by accounting for the proportion of liquid used and volume of air collected. All samples were corrected for the number of INPs on the average of four field blanks (cleaned, handled, transported, and analyzed in the same way without air flow). Two-tailed, 95% confidence intervals for binomial sampling are provided (Agresti and Coull, 1998). Select samples were also heat treated (95 °C for 20 min) to denature and deactivate biological INPs present and digested in 10% H 2 O 2 at 95 °C under UV-B for 20 min to remove any organic carbon INPs. Agresti, A, and BA Coull. 1998. "Approximate is better than “exact” for interval estimation of binomial proportions." American Statistics 52: 119–126. https://doi.org/10.2307/2685469 McCluskey, CS, J Ovadnevaite, M Rinaldi, J Atkinson, F Belosi, D Ceburnis, … and PJ DeMott. 2018. "Marine and Terrestrial Organic Ice-Nucleating Particles in Pristine Marine to Continentally Influenced Northeast Atlantic Air Masses." Journal of Geophysical Research: Atmospheres 123 (11): 6196–6212, https://doi.org/10.1029/2017JD028033 Vali, G. 1971. "Quantitative Evaluation of Experimental Results and the Heterogeneous Freezing Nucleation of Supercooled Liquids." Journal of the Atmospheric Sciences 28: 402-209. https://doi.org/10.1175/1520-0469(1971)028<0402:QEOERA>2.0.CO;2

54 ENVIRONMENTAL SCIENCES↗

DOE OSTI2026

Local Spin Density Approximation Strongly Improved by a Better-Informed Local Scaling of Its Self-Interaction Correction

The Perdew−Zunger self-interaction correction (PZSIC) makes density functional approximations (DFAs) exact for all one-electron densities. However, it overcorrects in manyelectron regions, introducing errors for the uniform-density limit, where uncorrected DFAs are exact. The locally scaled PZSIC (LSIC), based on the iso-orbital indicator zσ [which distinguishes single-orbital and slowly varying density regions and is used with the local spin density approximation (LSDA)], restores the uniform-density limit and significantly improves results for many properties, including chemical reaction barrier heights, atomization energies, and ionization potentials. Yet, LSIC performs poorly for weakly bonded systems, leaving many unbound, due to limitations of its iso-orbital indicator. To correct this, in this work we propose a new local scaling, LSIC-α, based on the iso-orbital indicator ασ (which additionally identifies regions of overlapping density tails). A two-parameter scaling function of ασ is fitted to a subset of the nonbonded appropriate norms for the SCAN and r2SCAN meta- GGAs, and tested on many properties of main-group atoms, molecules, and molecular complexes. LSIC-α greatly improves the interaction energies of weakly bonded systems in the S22 data set while retaining LSIC’s accuracy for other properties. This work shows that the errors of LSDA (and presumably of higher-level DFAs) can be largely but not entirely repaired by a proper “do no harm” self-interaction correction.

Approximation↗

NASA NTRS1995

Sailplane Glide Performance and Control Using Fixed and Articulating Winglets

An experimental study was conducted to investigate the effects of controllable articulating winglets on glide performance and yawing moments of high performance sailplanes. Testing was conducted in the Texas A&M University 7 x 10 foot Low Speed Wind Tunnel using a full-scale model of the outboard 5.6 feet of a 15 meter class high performance sailplane wing. Different wing tip configurations could be easily mounted to the wing model. A winglet was designed in which the cant and toe angles as well as a rudder on the winglet could be adjusted to a range of positions. Cant angles used in the investigation consisted of 5, 25, and 40 degrees measured from the vertical axis. Toe-out angles ranged from 0 to 22.5 degrees. A rudder on the winglet was used to study the effects of changing the camber of the winglet airfoil on wing performance and wing yawing moments. Rudder deflections consisted of-10, 0, and 10 degrees. Test results for a fixed geometry winglet and a standard wing tip are presented to show the general behavior of winglets on sailplane wings, and the effects of boundary-layer turbulators on the winglets are also presented. By tripping the laminar boundary-layer to turbulent before laminar separation occurs, the wing performance was increased at low Reynolds numbers. The effects on the lift and drag, yawing moment, pitching moment, and wing root bending moment of the model are presented. Oil flows were used on the wing model with the fixed geometry winglet and the standard wing tip to visualize flow directions and areas of boundary layer transition. A cant angle of 25 degrees and a toe-out angle of 2.5 degrees provided an optimal increase in wing performance for the cant and toe angles tested. Maximum performance was obtained when the winglet rudder remained in the neutral position of zero degrees. By varying the cant, toe, and rudder angles from their optimized positions, wing performance decreases. Although the winglet rudder proved to be more effective in increasing the yawing moment compared to varying the cant and toe angles, the amount of increased yawing moment was insignificant when compared to that produced by the vertical tail. A rudder on the winglet was determined to be ineffective for providing additional yaw control.

Colling, James David↗