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Models of two types of high-speed surface craft were tested in Langley tank no. 1 to obtain rough-water data for an evaluation by the David Taylor .Model Basin of the relative merits of the designs. Time-history records were obtained of trim, rise, and normal acceleration at two points in the hulls for various speeds and two sizes of waves.
An investigation of the effect of beam loading on impact loads and motions has been conducted in the Langley impact basin. Water impact tests of flat-bottom 5-inch- and 8-inch-beam models having beam-loading coefficients C(sub Delta) from 62.5 to 544 and a 30 0 dead-rise 5-inch-beam model A having beam-loading coefficients from 208 to 530 are described and the results analyzed to show trends of these heavy-beam-loading data with initial flight-path angle, trim angle, dead-rise angle, and time throughout the impact. Data from flat-bottom model tests, C(sub Delta) = 4.4 to 36.5, and from 300 dead-rise model tests, C(sub Delta)A = 0.58 and 18.8, are included, along with the heavy-beam-loading data; and variations of these data with beam-loading coefficients are shown. Each of the load and motion coefficients is found to be directly proportional to a power factor ofC(sub Delta). For instance, the maximum impact lift coefficient C(sub L,max)is found to be directly proportional to C(sub Delta)(sup 0.33) for the flat-bottom model and C(sub Delta)(sup 0.45) for the 30 deg dead-rise model. These variations of C(sub L,max) C(sub Delta) are found to be in agreement with theoretical variations. Finally, an empirical equation for the prediction of C(sub L,max) is presented and is shown to give good agreement with experimental C(sub L,max) for about 500 fixed-trim smooth-water impacts. The range of variables included dead-rise angles from 0 deg to 30 deg, beam-loading coefficients from 0.48 to 544, trim angles from 3 deg to 45 deg and initial flight-path angles from about 2 deg to about 27 deg.
The Tiltrotor Test Rig (TTR) is a new NASA asset for testing full-scale proprotors. The first test campaign in the National Full-Scale Aerodynamics Facility (NFAC) concluded in November 2018. The wind-tunnel test included vertical climb conditions simulated by axial flow at low airspeeds (true hover is not possible in the NFAC). The rotor tested was the Bell Model 699, an AW609 rotor modified specifically for wind-tunnel testing. The rotor was tested under a variety of NFAC configurations, some unprecedented and unique to vertical climb. This paper presents results for all test configurations applicable to vertical climb, including assessments of data quality. Comparisons with earlier tests of a similar rotor, the 0.656-scale Joint Vertical Experimental (JVX) rotor, are included to provide additional insights into rotor and wind tunnel behavior.
Data taken from tests at constant speed to establish trim limits of stability, tests at accelerated speeds to determine stable limits of center of gravity shift, and tests at decelerated speeds to obtain landing characteristics of several model hull forms were used to establish hull design effect on longitudinal stability of porpoising. Results show a reduction of dead rise angle as being the only investigated factor reducing low trim limit. Various methods of reducing afterbody interference increased upper trim limit
Polymer composites have properties such as high strength and stiffness, low weight, good thermal and chemical stability, as well as impact and abrasion resistance that make them ideal for high-performance applications. The chemistries of these materials are continuously improving, so determining their properties is vital for successfully producing them and achieving the desired results. Multiple methods can be employed to monitor characteristics such as heat flow, weight, dimension, and modulus as a function of time and temperature. By analyzing this information, models can be developed to predict outcomes of parameters not tested for. In one application, materials proposed for wet filament winding and the production of high pressure vessels can be analyzed to verify they will have the necessary low viscosity for good fiber wetting and long pot life for the extended handling inherent to this process. Such data about a prospective system provides valuable information on how that material could ultimately be processed to yield the desired part.
The Models for Aeroelastic Validation Research Involving Computation semi-span wind-tunnel model (MAVRIC-I), a business jet wing-fuselage flutter model, was tested in NASA Langley's Transonic Dynamics Tunnel with the goal of obtaining experimental data suitable for Computational Aeroelasticity code validation at transonic separation onset conditions. This research model is notable for its inexpensive construction and instrumentation installation procedures. Unsteady pressures and wing responses were obtained for three wingtip configurations clean, tipstore, and winglet. Traditional flutter boundaries were measured over the range of M = 0.6 to 0.9 and maps of Limit Cycle Oscillation (LCO) behavior were made in the range of M = 0.85 to 0.95. Effects of dynamic pressure and angle-of-attack were measured. Testing in both R134a heavy gas and air provided unique data on Reynolds number, transition effects, and the effect of speed of sound on LCO behavior. The data set provides excellent code validation test cases for the important class of flow conditions involving shock-induced transonic flow separation onset at low wing angles, including Limit Cycle Oscillation behavior.
The Models for Aeroelastic Validation Research Involving Computation semi-span wind-tunnel model (MAVRIC-I), a business jet wing-fuselage flutter model, was tested in NASA Langley's Transonic Dynamics Tunnel with the goal of obtaining experimental data suitable for Computational Aeroelasticity code validation at transonic separation onset conditions. This research model is notable for its inexpensive construction and instrumentation installation procedures. Unsteady pressures and wing responses were obtained for three wingtip configurations of clean, tipstore, and winglet. Traditional flutter boundaries were measured over the range of M = 0.6 to 0.9 and maps of Limit Cycle Oscillation (LCO) behavior were made in the range of M = 0.85 to 0.95. Effects of dynamic pressure and angle-of-attack were measured. Testing in both R134a heavy gas and air provided unique data on Reynolds number, transition effects, and the effect of speed of sound on LCO behavior. The data set provides excellent code validation test cases for the important class of flow conditions involving shock-induced transonic flow separation onset at low wing angles, including LCO behavior.
The supersonic aerodynamic and aeroelastic characteristics of a tension cone inflatable aerodynamic decelerator were investigated by wind tunnel testing. Two sets of tests were conducted: one using rigid models and another using textile models. Tests using rigid models were conducted over a Mach number range from 1.65 to 4.5 at angles of attack from -12 to 20 degrees. The axial, normal, and pitching moment coefficients were found to be insensitive to Mach number over the tested range. The axial force coefficient was nearly constant (C(sub A) = 1.45 +/- 0.05) with respect to angle of attack. Both the normal and pitching moment coefficients were nearly linear with respect to angle of attack. The pitching moment coefficient showed the model to be statically stable about the reference point. Schlieren images and video showed a detached bow shock with no evidence of large regions of separated flow and/or embedded shocks at all Mach numbers investigated. Qualitatively similar static aerodynamic coefficient and flow visualization results were obtained using textile models at a Mach number of 2.5. Using inflatable textile models the torus pressure required to maintain the model in the fully-inflated configuration was determined. This pressure was found to be sensitive to details in the structural configuration of the inflatable models. Additional tests included surface pressure measurements on rigid models and deployment and inflation tests with inflatable models.
The Sublimator Driven Coldplate (SDC) is a unique piece of thermal control hardware that has several advantages over a more traditional thermal control system. The principal advantage is the possible elimination of a pumped fluid loop, potentially saving mass, power, and complexity. Because this concept relies on evaporative heat rejection techniques, it is primarily useful for short mission durations. Additionally, the concept requires a conductive path between the heat-generating component and the heat rejection device. Therefore, it is mostly a relevant solution for a vehicle with a relatively low heat rejection requirement and/or short transport distances. Tests were performed on coupons and an Engineering Development Unit (EDU) at NASA s Johnson Space Center to better understand the basic operational principles and to validate the analytical methods being used for the SDC development. This paper outlines the results of the SDC tests, the subsequent thermal model correlation, and a description of the SDC Engineering Development Unit test results.
The concept employed in an existing Trajectory Control Sensor (TCS) breadboard is being developed into an engineering model to be considered for flight on the Shuttle as a Detailed Test Objective (DTO). The sensor design addresses the needs of Shuttle/SSF docking/berthing by providing relative range and range rate to 1500 meters as well as the perceived needs of AR&C by relative attitude measurement over the last 100 meters. Range measurement is determined using a four-tone ranging technique. The Doppler shift on the highest frequency tone will be used to provide direct measurement of range rate. Bearing rate and attitude rates will be determined through back differencing of bearing and attitude, respectively. The target consists of an isosceles triangle configuration of three optical retroreflectors, roughly one meter and one-half meter in size. After target acquisition, the sensor continually updates the positions of the three retros at a rate of about one hertz. The engineering model is expected to weigh about 25 pounds, consume 25-30 watts, and have an envelope of about 1.25 cubic feet. The following concerns were addressed during the presentation: are there any concerns with differentiating attitude and bearing to get attitude and bearing rates? Since the docking scenario has low data bandwidth, back differencing is a sufficient approximation of a perfect differentiator for this application. Could range data be obtained if there were no retroreflectors on the target vehicle? Possibly, but only at close range. It would be dependent on target characteristics.
Intelligent software is being developed for closed life support systems with biological components, for human exploration of the Moon and Mars. The intelligent software functions include planning/scheduling, reactive discrete control and sequencing, management of continuous control, and fault detection, diagnosis, and management of failures and errors. Four types of modeling information have been essential to system modeling and simulation to develop and test the software and to provide operational model-based what-if analyses: discrete component operational and failure modes; continuous dynamic performance within component modes, modeled qualitatively or quantitatively; configuration of flows and power among components in the system; and operations activities and scenarios. CONFIG, a multi-purpose discrete event simulation tool that integrates all four types of models for use throughout the engineering and operations life cycle, has been used to model components and systems involved in the production and transfer of oxygen and carbon dioxide in a plant-growth chamber and between that chamber and a habitation chamber with physicochemical systems for gas processing.
An investigation of the effect of beam loading on impact loads and motions has been conducted in the Langley impact basin. Water impact tests of flat-bottom 5-inch-and 8-inch-beam models having beam-loading coefficients C(sub Delta) from 62.5 to 544 and a 30 deg dead-rise 5-inch-beam model A having beam-loading coefficients from 208 to 530 are described and the results analyzed to show trends of these heavy-beam-loading data with initial flight-path angle, trim angle, dead-rise angle, and time through- out the impact. Data from flat-bottom model tests, C(sub Delta) = 4.4 to 36.5, and from 30 deg dead-rise model tests, C(sub Delta) = 0.58 and 18.8, are included, along with the heavy-beam-loading data; and variations of these data with beam-loading coefficients are shown. Each of the load and motion coefficients is found to be directly proportional to a power factor of C(sub Delta). For instance, the maximum impact lift coefficient is found C(sub L,max) to be directly proportional to C(sub Delta, exp 0,33) for the flat-bottom model and C(sub Delta, exp 0,45) for the 30 deg dead-rise model. These variations of C(sub L, max) with C(sub Delta) are found to be in agreement with theoretical variations. Finally, an empirical equation for the prediction of C(sub L,max) is presented and is shown to give good agreement with experimental C(sub L, max) for about 500 fixed-trim smooth-water impacts. The range of variables included dead-rise angles from 0 deg to deg, beam-loading coefficients from o.48 to 544, trim angles from 3 deg to deg, and initial flight-path angles from about 2 deg to about deg.
Experimental studies to determine the 140-deg spherically-blunted cone configuration single degree-of-freedom damping derivative included model sting interference effects at M = 1.76 to 3.0. Minimal sting influences are noted with sting/model diameter ratios of 0.183 and effective sting lengths of 3.55 of the reference length. Increasing the ratio of sting diameter to reference length to 0.53 influences damping considerably at angles of attack from 0 to 3 deg at all Reynolds numbers but shows little or no effect at higher angles. Free-flight model wake geometries (wake neck position and size) correlate well with those from captive model tests with the small diameter sting; this adds confidence to obtaining damping derivatives through captive model testing. The damping characteristics of an early Viking configuration obtained on 11.23 and 4.35% scale models using forced and free oscillation test methods, respectively, in two different test facilities are compared. The agreement is quite good under test conditions.
This project explores electrolytic hydrogen production from an offshore wind turbine by: Modeling: Simulate an offshore wind turbine and generate power output profiles; Analysis: Analyze offshore conditions and determine operational modes of the stack; Testing: Perform hardware testing to evaluate dynamic characteristics; and International Collaboration Work: Share learnings and publish. Potential Outcomes: Accelerate development of an integrated, in-turbine offshore wind hydrogen system; and Support two DOE Energy Earthshots - Hydrogen and Floating Offshore Wind.
Here, we present a one-dimensional (1-D) model composed of aligned, electrostatically interacting charged disks, conceived to address in a computable model the validity of the Bogoliubov assumption on the decay of particle correlations in the Born–Bogoliubov–Green–Kirkwood–Yvon hierarchy. This assumption is a basic premise of plasma kinetic theory. The disk model exhibits spatially 1-D features at short distances, but retains 3-D features at large distances. Here the collective dynamics of this model plasma is investigated by solving the corresponding Vlasov equation. In addition, the implementation of the model for the numerical validation of the Bogoliubov assumption is formulated.
The Rhesus monkey has been proposed as a model for the effects of space flight on immunity. In order to determine the feasibility of the use of the Rhesus monkey as a model, we studied the use of Rhesus monkey cells for immunological procedures that have been shown to be affected by space flight in both rodents and humans. We have shown that both lymph node cells and peripheral blood leukocytes can be stained with monoclonal antibodies to detect the following surface markers: CD4, CD-8, Ia and surface immunoglobulin. Also, the level of Ia antigen expression was increased by treatment of the cells with human interferon-gamma. In addition, cells were induced to produce interferons and interleukins. Isolated neutrophils also demonstrated increased oxidative burst. These data indicate that the Rhesus monkey will be a useful model for space flight studies of immunity.
The topics covered include the following: 20 dB jet noise suppression; ejector nozzle technology program - noise reduction vs. flow augmentation; mixer ejector nozzle technology challenges; 1989 High Speed Civil Transport (HSCT) 2-D ejector model test in NASA's 9 x 15 ft. tunnel; tertiary airflow 1989 2-D ejector test; shock noise dominates 2-D ejector test; lessons learned - 2-D mixer/ejector in 9 x 15 ft. tunnel; 1990 HSCT axisymmetric ejector model test in Boeing's Low Speed Aeroacoustic Facility (LSAF); axisymmetric mixer/ejector mach contours - peak and valley - NASTAR pre-test predictions; tertiary airflow objectives accomplished - 1990 AXI model; and HSCT low noise exhaust technology programs.
The differential equation for the trajectory of a spherical particle injected into an aircraft wake was developed and the proper scaling relations extracted. After some simplification a convenient set of similarity parameters was established. Using these similarity parameters a scale model test program was designed and performed in the NASA Langley vortex research facility. The results of the tests demonstrated the validity of the similarity parameters in conducting scale model testing for aerial application research.