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Brief discussions of the materials and configurations testing techniques and requirements for the Apollo spacecraft high pressure oxygen systems are presented.
An investigation was conducted in the Langley 30- by 60-Foot Tunnel to determine the aerodynamic characteristics of a powered, full-scale model of a general aviation airplane employing a canard. Although primary emphasis of the investigation was placed on evaluating the aerodynamic performance and the stability and control characteristics of the basic configuration, tests were also conducted to study the following effects of varying the basic configuration: effect of Reynolds number; effect of canard; effect of outboard wing leading-edge droop; effect of center-of-gravity location; effect of elevator trim; effect of landing gear; effect of lateral-directional control; effect of power; effect of fixed transition; effect of water spray; effects of canard incidence, canard airfoil section, and canard position; and effects of winglets and upper winglet size. Additional aspects of the study were to determine the boundary-layer transition characteristics of airfoil surfaces and the effect of fixing the boundary layer to be turbulent by means of a transition strip near the leading edge. The tests were conducted at Reynolds numbers from 0.60 x 10 to the 6th power to 2.25x10 to the 6th power, based on the wing mean aerodynamic chord, at angles of attack from -4.5 deg to 41.5 deg, and at angles of sideslip from -15 deg to 15 deg.
The limitations of structural design with 3D woven composites based on homogenized material models characterized using coupon test data, representative unit cell analysis, or a combination of the two are discussed in the context of a space launch system (SLS)-scale end ring. Selected building block test configurations at the coupon, element, and subcomponent scales are described in detail. Highly instrumented tests and X-ray computed tomography inspections are used to characterize the structural behavior and associated damage progression with increasing load. It is shown that a single homogenized material constitutive relationship, applied at coupon and element scale, is not sufficient to model the structural response observed in both tests. In fact, substantially different stress-strain laws are required to predict the response of the coupon and element scale specimens, despite similar loading conditions. A model with discrete representation of the binder tows and cohesive interfaces to model matrix cracking shows much better agreement with the test data at the two length scales. Thus, evidence is provided showing that predictive structural analysis of 3D woven composites requires discrete representation of the tows, matrix cracking morphology, and their interactions in the structural configuration. With accurate knowledge of the element scale behavior, engineering approximations readily enable analysis scale-up to the subcomponent level.
Device durability remains a significant challenge in photoelectrochemical (PEC) water splitting under ambient conditions. Yet, a lack of understanding of the test configuration and applied bias effects continue to hinder progress. In this study, we differentiate two-electrode (2E) and three-electrode (3E) configurations for evaluating PEC material durability, focusing particularly on their impacts on photoabsorber solid-state operating conditions. Our results underscore the fallacy of inferring 2E device stability from durability measurements performed solely in 3E configurations. Unmeasured and often misunderstood total circuit bias in 3E tests moderates material degradation, leading to the overestimation of photoelectrode stability compared to short-circuit operation. We demonstrate how the photoabsorber's operating voltage critically governs charge separation, surface stability, and degradation mechanisms during PEC operation. With these findings, we propose a standardized framework for conducting more reliable 3E durability experiments that simulate unassisted performance to help accelerate the development of robust, stable materials for solar-driven water splitting.
A limited experimental investigation was conducted in the Langley 4- by 7-Meter Tunnel to explore the effects of aft-fuselage-mounted advanced turboprop installations on the low-speed stability and control characteristics of a representative transport aircraft in a landing configuration. In general, the experimental results indicate that the longitudinal and lateral-directional stability characteristics for the aft-fuselage-mounted single-rotation tractor and counter-rotation pusher propeller configurations tested during this investigation are acceptable aerodynamically. For the single-rotation tractor configuration, the propeller-induced aerodynamics are significantly influenced by the interaction of the propeller slipstream with the pylon and nacelle. The stability characteristics for the counter-rotation pusher configuration are strongly influenced by propeller normal forces. The longitudinal and directional control effectiveness, engine-out characteristics, and ground effects are also presented. In addition, a tabulated presentation of all aerodynamic data presented in this report is included as an appendix.
The experimental/analytical research work described here addresses the rocket-ejector mode (Mach 0-2 operational range) of the RBCC engine. The experimental phase of the program includes studying the mixing and combustion characteristics of the rocket-ejector system utilizing state-of-the-art diagnostic techniques. A two-dimensional variable geometry rocket-ejector system with enhanced optical access was utilized as the experimental platform. The goals of the experimental phase of the research being conducted at Penn State are to: (a) systematically increase the range of rocket-ejector understanding over a wide range of flow/geometry parameters and (b) provide a comprehensive data base for evaluating and anchoring CFD codes. Concurrent with the experimental activities, a CFD code benchmarking effort at Marshall Space Flight Center is also being used to further investigate the RBCC rocket-ejector mode. Experiments involving the single rocket based optically-accessible rocket-ejector system have been conducted for Diffusion and Afterburning (DAB) as well as Simultaneous Mixing and Combustion configurations. For the DAB configuration, air is introduced (direct-connect) or ejected (sea-level static) into a constant area mixer section with a centrally located gaseous oxygen (GO2)/gaseous hydrogen (GH2) rocket combustor. The downstream flowpath for this configuration includes a diffuser, an afterburner and a final converging nozzle. For the SMC configuration, the rocket is centrally located in a slightly divergent duct. For all tested configurations, global measurements of the axial pressure and heat transfer profiles as well as the overall engine thrust were made. Detailed measurements include major species concentration (H2 O2 N2 and H2O) profiles at various mixer locations made using Raman spectroscopy. Complementary CFD calculations of the flowfield at the experimental conditions also provide additional information on the physics of the problem. These calculations are being conducted at Marshall Space Flight Center to benchmark the FDNS code for RBCC engine operations for such configurations. The primary fluid physics of interests are the mixing and interaction of the rocket plume and secondary flow, subsequent combustion of the fuel rich rocket exhaust with the secondary flow and combustion of the injected afterburner flow. The CFD results are compared to static pressure along the RBCC duct walls, Raman Spectroscopy specie distribution data at several axial locations, net engine thrust and entrained air for the SLS cases. The CFD results compare reasonably well with the experimental results.
In January 2004, shortly after the Columbia accident, future servicing missions to the Hubble Space Telescope (HST) were cancelled. In response to this, further work on the Wide Field Camera 3 instrument was ceased. Given the maturity level of the design, a characterization thermal test (TV1) was completed in case the mission was re-instated or an alternate mission found on which to fly the instrument. This thermal test yielded some valuable lessons learned with respect to testing configurations and modeling/correlation practices, including: 1. Ensure that the thermal design can be tested 2. Ensure that the model has sufficient detail for accurate predictions 3. Ensure that the power associated with all active control devices is predicted 4. Avoid unit changes for existing models. This paper documents the difficulties presented when these recommendations were not followed.
Ares I-X was the first flight test vehicle used in the development of NASA's Ares I crew launch vehicle. The Ares I-X used a 4-segment reusable solid rocket booster from the Space Shuttle heritage with mass simulators for the 5th segment, upper stage, crew module and launch abort system. Three modal tests were defined to verify the dynamic finite element model of the Ares I-X flight test vehicle. Test configurations included two partial stacks and the full Ares I-X flight test vehicle on the Mobile Launcher Platform. This report focuses on the first modal test that was performed on the top section of the vehicle referred to as Stack 5, which consisted of the spacecraft adapter, service module, crew module and launch abort system simulators. This report describes the test requirements, constraints, pre-test analysis, test operations and data analysis for the Ares I-X Stack 5 modal test.
Ares I-X was the first flight test vehicle used in the development of NASA s Ares I crew launch vehicle. The Ares I-X used a 4-segment reusable solid rocket booster from the Space Shuttle heritage with mass simulators for the 5th segment, upper stage, crew module and launch abort system. Three modal tests were defined to verify the dynamic finite element model of the Ares I-X flight test vehicle. Test configurations included two partial stacks and the full Ares I-X flight test vehicle on the Mobile Launcher Platform. This report focuses on the second modal test that was performed on the middle section of the vehicle referred to as Stack 1, which consisted of the subassembly from the 5th segment simulator through the interstage. This report describes the test requirements, constraints, pre-test analysis, test operations and data analysis for the Ares I-X Stack 1 modal test.
A software-based thermal modeling process was documented for generating the thermal panel settings necessary to simulate worst-case on-orbit flight environments in an observatory-level thermal vacuum test setup. The method for creating such a thermal model involved four major steps: (1) determining the major thermal zones for test as indicated by the major dissipating components on the spacecraft, then mapping the major heat flows between these components; (2) finding the flight equivalent sink temperatures for these test thermal zones; (3) determining the thermal test ground support equipment (GSE) design and initial thermal panel settings based on the equivalent sink temperatures; and (4) adjusting the panel settings in the test model to match heat flows and temperatures with the flight model. The observatory test thermal model developed from this process allows quick predictions of the performance of the thermal vacuum test design. In this work, the method described above was applied to the Global Precipitation Measurement (GPM) core observatory spacecraft, a joint project between NASA and the Japanese Aerospace Exploration Agency (JAXA) which is currently being integrated at NASA Goddard Space Flight Center for launch in Early 2014. From preliminary results, the thermal test model generated from this process shows that the heat flows and temperatures match fairly well with the flight thermal model, indicating that the test model can simulate fairly accurately the conditions on-orbit. However, further analysis is needed to determine the best test configuration possible to validate the GPM thermal design before the start of environmental testing later this year. Also, while this analysis method has been applied solely to GPM, it should be emphasized that the same process can be applied to any mission to develop an effective test setup and panel settings which accurately simulate on-orbit thermal environments.
The X-56A Multi-Utility Technology Testbed (MUTT) is a subscale, fixed-wing aircraft designed for high-risk aeroelastic flight demonstration and research. Structural dynamics ground testing for model validation was especially important for this vehicle because the structural model was directly used in the development of a flight control system with active flutter suppression capabilities. Structural dynamics ground tests of the X-56A MUTT with coupled rigid-body and structural modes provided a unique set of challenges. An overview of the ground vibration test (GVT) and moment of inertia (MOI) test setup and execution is presented. The series of GVTs included the wing by itself attached to a strongback and complete vehicle at two mass conditions: empty and full fuel. Two boundary conditions for the complete-vehicle test were studied: on landing gear and suspended free-free. Pitch MOI tests were performed using a compound pendulum method and repeated with two different pendulum lengths for independent verification. The original soft-support test configuration for the GVT used multiple bungees, resulting in unforeseen coupling interactions between the soft-support bungees and the vehicle structural modes. To resolve this problem, the soft-support test setup underwent multiple iterations. The various GVT configurations and boundary-condition modifications are highlighted and explained. Lessons learned are captured for future consideration when performing structural dynamics testing with similar vehicles.
Tests were conducted with plain bearings running with three types of three-lobe journals. The bearings, with a nominal 3.8-centimeter (1.5-in.) diameter and L/D of 1, were tested at 300 K (80 F) under stable conditions to 5400 rpm. The lobed configurations tested, in order of diminishing stability, were a tilted-lobe journal with grooves, a tilted-lobe journal without grooves, and a centrally lobed journal without grooves. A previously tested three-tilted-lobe bearing with grooves, running with a plain journal, was somewhat more stable than the tilted-lobe or centrally lobed journals.
Data obtained from a heat transfer test conducted on an 0.006-scale space shuttle orbiter and external tank in the NASA-Ames Research Center 3.5-foot Hypersonic Wind Tunnel are presented. The purpose of this test was to obtain data under simulated return-to-launch-site abort conditions. Configurations tested were integrated orbiter and external tank, orbiter alone, and external tank alone at angles of attack of 0, + or - 30, + or - 60, + or - 90, and + or - 120 degrees. Runs were conducted at Mach numbers of 5.2 and 5.3 for Reynolds numbers of 1.0 and 4.0 million per foot, respectively. Heat transfer data were obtained from 75 orbiter and 75 external tank iron-constantan thermocouples.
Emissions data were obtained for six fuel injector configurations tested with ultra lean combustion. Fuel injectors included three multiple source designs and three configurations using a single air assist injector. Only the multiple source fuel injectors provided acceptable emissions. Values of 16g CO/kg fuel, 1.9g HC/kg fuel, and 19.g NO2/kg fuel were obtained for the combustion temperature range of 1450 to 1700 K for both a high blockage 19 source injector and a low blockage 41 source injector. It was shown that high fuel injector pressure drop may not be required to achieve low emissions performance at high inlet air temperature when the fuel is well dispersed in the airstream.
A detailed discussion of Launch Processing System Ground Software Production is presented to establish the interrelationships of firing room resource utilization, configuration control, system build operations, and Shuttle data bank management. The production of a test configuration identifier is traced from requirement generation to program development. The challenge of the operational era is to implement fully automated utilities to interface with a resident system build requirements document to eliminate all manual intervention in the system build operations. Automatic update/processing of Shuttle data tapes will enhance operations during multi-flow processing.
Proof-of-concept testing was performed on a 20-kHz, resonant power system breadboard from 1981 through 1985. The testing began with the evaluation of a single, 1.0-kW resonant inverter and progressed to the testing of breadboard systems with higher power levels and more capability. The final breadboard configuration tested was a 25.0-kW breadboard with six inverters providing power to three user-interface modules over a 50-meter, 20-kHz bus. The breadboard demonstrated the ability to synchronize multiple resonant inverters to power a common bus. Single-phase and three-phase 20-kHz power distribution was demonstrated. Simple conversion of 20-kHz to dc and variable-frequency ac was demonstrated as was bidirectional power flow between 20-kHz and dc. Steady state measurements of efficiency, power-factor tolerance, and conducted emissions and conducted susceptibility were made. In addition, transient responses were recorded for such conditions as start up, shut down, load changes. The results showed the 20-kHz resonant system to be a desirable technology for a spacecraft power management and distribution system with multiple users and a utility-type bus.
A validation procedure for the Ada binding of the Graphical Kernel System (GKS) is being developed. PRIOR Data Sciences is also producing a version of the GKS written in Ada. These major software engineering projects will provide an opportunity to demonstrate a sound approach for software testing in an Ada environment. The GKS/Ada validation capability will be a collection of test programs and data, and test management guidelines. These products will be used to assess the correctness, completeness, and efficiency of any GKS/Ada implementation. The GKS/Ada developers will be able to obtain the validation software for their own use. It is anticipated that this validation software will eventually be taken over by an independent standards body to provide objective assessments of GKS/Ada implementations, using an approach similar to the validation testing currently applied to Ada compilers. In the meantime, if requested, this validation software will be used to assess GKS/Ada products. The second project, implementation of GKS using the Ada language, is a conventional software engineering tasks. It represents a large body of Ada code and has some interesting testing problems associated with automatic testing of graphics routines. Here the normal test practices which include automated regression testing, independent quality assistance, test configuration management, and the application of software quality metrics will be employed. The software testing methods emphasize quality enhancement and automated procedures. Ada makes some aspects of testing easier, and introduces some concerns. These issues are addressed.
Nonlinear-tapered flexbeam specimens from a full-size composite helicopter rotor hub flexbeam were tested under combined constant axial tension and cyclic bending loads. Two different graphite/glass hybrid configurations tested under cyclic loading failed by delamination in the tapered region. A 2-D finite element model was developed which closely approximated the flexbeam geometry, boundary conditions, and loading. The analysis results from two geometrically nonlinear finite element codes, ANSYS and ABAQUS, are presented and compared. Strain energy release rates (G) associated with simulated delamination growth in the flexbeams are presented from both codes. These results compare well with each other and suggest that the initial delamination growth from the tip of the ply-drop toward the thick region of the flexbeam is strongly mode II. The peak calculated G values were used with material characterization data to calculate fatigue life curves for comparison with test data. A curve relating maximum surface strain to number of loading cycles at delamination onset compared well with the test results.