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Nonlinear Slosh Damping Testing and Analysis for Launch Vehicle Propellant Tanks
Propellant tank slosh damping models for launch vehicle ascent analysis typically employ linear models for bare-wall damping, or assume a single, low-amplitude wave height for baffled configurations. A higher fidelity damping model would incorporate nonlinear effects to increase damping as slosh wave amplitude increases. This paper provides an overview of a technical assessment performed by the NASA Engineering and Safety Center (NESC) to evaluate lateral slosh damping as a function of lateral force (or wave) amplitude for multiple tank configurations. The increased fidelity of nonlinear slosh damping models, if leveraged, has the potential to reduce over-conservatism associated with the use of linear slosh damping models. This can provide design flexibility to enhance launch vehicle flight control performance, reduce baffle design requirements and/or increase robustness in targeted areas such as control-structure interaction during ascent.
Testing of TVS Augmented Injectors for Tank-to-Tank Transfer of Cryogens
NASA is developing reusable cryogenic systems to enable manned missions to the Moon and Mars. These cryogenic systems will require replenishing cryogens on-orbit via a tanker or propellant depot with minimal loss to boil-off. Three additively manufactured injectors augmented with a Thermodynamic Venting System have been developed for use in tank-to-tank cryogenic fluid transfer. A Vented Chill/No-Vent Fill approach to tank-to-tank transfer has been the industry standard and is explored in this work with emphasis on injector spray patterns and compared to the Charge-Hold-Vent chill method.
Testing of TVS Augmented Injectors for Tank-to-Tank Transfer of Cryogens
NASA is developing reusable cryogenic systems to enable manned missions to the Moon and Mars. These cryogenic systems will require replenishing cryogens on-orbit via a tanker or propellant depot with minimal loss to boil-off. Three additively manufactured injectors augmented with a Thermodynamic Venting System have been developed for use in tank-to-tank cryogenic fluid transfer. A Vented Chill/No-Vent Fill approach to tank-to-tank transfer has been the industry standard and is explored in this work with emphasis on injector spray patterns and compared to the Charge-Hold-Vent chill method.
Results of variable enthalpy tests of CPR-488 TIP panels in MSFC Hot Gas Facility
Tests were conducted to determine the effects of variable enthalpy levels on CPR-488 foam performance and/or recession. Also to be determined was whether the BX-250 thickness indicator plugs (TIP) concept is feasible for measuring recession rates of CPR-488 foam during external tank flight. The tests determined that the TIP plug idea works. The accuracy of the indicated recession depth is approximately + or - 1.16 in. The BX-250 plugs with TiO2 work better than the regular BX-250 plugs. Also, there is an effect of enthalpy on CPR-488 foam performance.
Shuttle ascent and shock impingement aerodynamic heating studies
The collection and analysis of aerodynamic heating data obtained from shock impingement experimental investigation were completed. The data were categorized into four interference areas; fin leading edge, wing/fuselage fin/plate corners, and space shuttle configurations. The effects of shock impingement were found to increase the heating rates 10 to 40 times the undisturbed values. A test program was completed at NASA/Langley Research Center to investigate the magnitudes and surface patterns of the mated shock interference flowfield. A 0.0065 scale thin-skin model of the MDAC 256-20 space shuttle booster mated with a Stycast model of the MDAC Internal tank orbiter was tested in the 20-inch M=6 tunnel, the 31-inch M=10 tunnel, and the 48-inch Unitary Plan Tunnel. The gap region of the ascent configuration was the principal area of interest where both thermocouple and phase-change paint data were obtained. Pressure and heat transfer distributions data on the leeward surface of a 75-degree sweep slab delta wing are presented. The effects of surface roughness on boundary layer transition and aerodynamic heating were investigated.
Hydrodynamic Tests of a 1/10-size Model of the Hull of the Latecoere 521 Flying Boat : NACA Model 83
A 1/10-size model of the hull of the French flying boat Latecoere 521 was tested in the NACA tank. This model is one of a series of models of the hulls of actual flying boats of both foreign and domestic type that are being tested in the NACA tank to provide information regarding the water characteristics of a variety of forms of hull and to illustrate the development of present-day types of flying boat. The lines and the offsets of the hull were obtained from the manufacturer through the Paris office of the NACA. The form of the stub-wing stabilizers was not furnished and, therefore, the model was tested without them. The model was tested free to trim at the design initial load (initial load coefficient of 0.428) and by the general method at load coefficients from 0.025 to 0.6. The spray characteristics of the model are good. The form of the bow would be particularly desirable for rough water use. The interference of the afterbody and the tail extension is excessive, causing very high resistance at high speeds. A violent vertical instability is present at trims of 4 degrees and 6 degrees with light loads and high speeds.
Spin Investigation of a 1/29-Scale Model of the Republic XF-91 Airplane with a Conventional Tail Installed
An investigation has been conducted in the Langley 20-foot free-spinning tunnel of a 1/29-scale model of the Republic XF-91 airplane with a.conventional-tail arrangement installed. Previously, tests were made on the model with a vee tail installed. The erect spin and recovery characteristics of the model were determined for the normal loading with the wing installed at various amounts of incidence. The spin investigation also included inverted-spin tests, spin-recovery-parachute tests, tests with the center of gravity moved rearward, and tests with external fuel tanks added to the model. In addition, several tail.modifications were tested,on the model in an attempt, to improve the model's spin-recovery characteristics. The results indicate that any fully developed spin obtained on the airplane with the conventional tail installed will be satisfactorily terminated if rudder reversal is accompanied by moving the ailerons with the spin (stick right in a right spin).Decreasing the wing incidence from 6deg to -2deg should have a beneficial effect on the recovery characteristics of the airplane. Recovery characteristics by normal use of controls (full rudder reversal followed by moving the elevators down) will be satisfactory if the wing incidence,of the airplane is -2deg. Installation of external fuel tanks (with or without fuel) will have a somewhat adverse effect on the recovery characteristics of the airplane, but if the recovery technique includes movement of the ailerons to full with the spin, the spin rotation will be terminated rapidly. Varying the position of the center of gravity within the limits indicated to be possible on the airplane should not affect the recovery characteristics.
SS/RCS surface tension propellant acquisition/expulsion tankage technology
The analysis, design, fabrication, and testing of a propellant tank that satisfies the requirements of the space shuttle is presented. This mission presents very stringent and sometimes conflicting requirements. A compartmented-tank device was developed and various ground and drop tower test techniques were employed to verify the design using both subscale and full-scale hardware. Performance was established with scale models and further substantiation was obtained with the full-scale tankage. Fabrication, acceptance, fill and drain, inspection, and other ground handling procedures were developed.
NDE Methods for the External Tank Spray on Foam Insulation
The foam thermal protection system (TPS) of the space shuttle External Tank (ET) has provided some unique challenges to the nondestructive testing community. Three nondestructive evaluation methods have been developed to identify defects in the foam TPS of the ET. Terahertz imaging and backscatter radiography have been developed to identify voids in thick foam regions while shearography has been developed to identify shallow delaminations, shallow voids and crush damage in the foam. The basic theory of operation along with factors, determined from a series of designed experiments, affecting the results of these methods will be described. Results from both test panels and flight tank inspections will be provided, along with predicted probability of detection estimates to show the range in defect sizes and types that can be readily detected.
STS-102 Countdown Status Briefing
Joel Wells, NASA Public Affairs, introduces Pete Nickolenko, NASA Test Director, Glenn Chin, Leonardo Mission Manager, and Ed Priselac, Shuttle Weather Officer, in this STS-102 prelaunch press conference. Mr. Nickolenko gives an overview of the countdown and built-in hold times, the launch window, and prelaunch activities (such as activation and checkout of the onboard computer systems, closing the payload bay doors, servicing of the onboard cryogenic cell tanks, main engine tests, and power-up of the ground communications systems). Mr. Chin confirms that the payload is in the final flight configuration and is ready for launch. Mr. Priselac gives the weather forecast for the launch date. The men then answer questions from the press.
Hypersonic aerodynamic characteristics of NR-ATP orbiter, orbiter with external tank, and ascent configuration
A scale model of the North American Rockwell ATP Orbiter with and without the external tank has been tested in a 22-inch helium tunnel at Mach 20 and a Reynolds number based on model length, of 2.14 times one million. Longitudinal and lateral-directional data were determined for the orbiter alone while only longitudinal characteristics and elevon roll effectiveness were investigated for the orbiter/tank combination. Oil flow and electron beam flow visualization studies were conducted for the orbiter alone, orbiter with external tank and the ascent configuration.
Yawed-Landing Investigation of a Model of the Convair Y2-2 Airplane, TED No. NACA DE 363
A model of the Convair Y2-2 airplane was tested in Langley tank no. 2 to determine whether satisfactory stability in yawed landings was possible with a certain ventral fin. Free-body landings were made in smooth and rough water at two speeds and two rates of descent with the model yawed 15deg. The behavior of the model was determined by visual observations and from motion-picture re.cords. It was concluded that satisfactory stability was possible with the ventral fin as tested but that the characteristics of the model shock absorbers and the settings of the elevon control surfaces had an appreciable influence on behavior.
Fracture Testing of Integral Stiffened Structure
Laboratory testing was conducted to evaluate safety concerns for integrally-stiffened tanks that were found to have developed cracks during pressurization testing. Cracks occurred at fastener holes where additional stiffeners were attached to the integrally-stiffened tank structure. Tests were conducted to obtain material properties and to reproduce the crack morphologies that were observed in service to help determine if the tanks are safe for operation. Reproducing the cracking modes observed during pressurization testing required a complex loading state involving both a tensile load in the integrally-stiffened structure and a pin-load at a fastener hole.
Results of oil flow visualization tests of an 0.010-scale model (52-OT) of the space shuttle orbiter-tank mated and orbiter configurations in the AEDC VKF tunnel B (IA17B)
An 0.010-scale model of the space shuttle (orbiter-tank mated and orbiter configurations) was tested in the AEDC VKF Tunnel B to investigate aerodynamic flow patterns. The tests utilized oil flow techniques to visualize the flow patterns. Tunnel free stream Mach number was 7.95 and nominal unit Reynolds number was 3.7 million per foot. Model angle of attack was varied from -5 deg through 10 deg and angle of sideslip was 0 deg and 2 deg. Photographs of resulting oil flow patterns are presented.
Inspection of the Space Shuttle External Tank SOFI Using Near-Field and Focused Millimeter Wave Nondestructive Testing Techniques
The Space Shuttle Columbia's catastrophic failure has been attributed to a piece of external tank SOFI (Spray On Foam Insulation) striking the left wing of the orbiter causing significant damage to some of the reinforced carbon/carbon leading edge wing panels. Subsequently, several nondestructive testing (NDT) techniques have been considered for inspecting the external tank. One such method involves using millimeter waves which have been shown to easily penetrate through the foam and provide high resolution images of its interior structures. This paper presents the results of inspecting three different SOFI covered panels by reflectometers at millimeter wave frequencies, specifically at 100 GHz. Each panel was fitted with various embedded anomalies/inserts representing voids and unbonds of diferent shapes, sizes and locations within each panel. In conjunction with these reJqectome&rs, radiators including a focused lens antenna and a small horn antenna were used. The focused lens antenna provided for a footprint diameter of approximately 1.25 cm (0.5") at 25.4 cm (10") away from the lens surface. The horn antenna was primarily operated in its near-field for obtaining relatively high resolution images. These images were produced using 2 0 scanning mechanisms. Discussions of the difference between the capabilities of these two types of antennas (radiators) for the purpose of inspecting the SOFI as it relates to the produced images are also presented.
TPSAS-NF1676L-13579-DND
During aerospace vehicle conceptual and preliminary design, empirical non-optimum factors are typically applied to predicted structural component weights to account for undefined manufacturing and design details. Non-optimum factors are developed here for 32 aluminum-lithium 2195 orthogrid panels comprising the liquid hydrogen tank barrel of the Space Shuttle External Tank using measured panel weights and manufacturing drawings. Minimum values for skin thickness, axial and circumferential blade stiffener thickness and spacing, and overall panel thickness are used to estimate individual panel weights. Panel non-optimum factors computed using a coarse weights model range from 1.21 to 1.77, and a refined weights model (including weld lands and skin and stiffener transition details) yields non-optimum factors of between 1.02 and 1.54. Acreage panels have an average 1.24 non-optimum factor using the coarse model, and 1.03 with the refined version. Non-optimum factors are also calculated for the aluminum-alloy gores that compose the ogive and ellipsoidal bulkheads of the Space Shuttle Super-Lightweight Tank propellant tanks. Minimum values for actual gore skin thicknesses and weld land dimensions are extracted from component production drawings, and are used to predict reference gore weights. These actual skin thicknesses are also compared to skin thicknesses predicted using classical structural mechanics and tank proof-test pressures. Both coarse and refined weights models are developed for the gores. The coarse model is based on the proof-test skin thicknesses, and the refined model uses the actual skin thicknesses and weld land details. Both models include manufacturing tolerance weight estimates. To determine the gore non-optimum factors, these reference weights are then compared to flight hardware weights reported in a mass properties database. The gore non-optimum factors computed using the coarse weights model range from 1.18 to 2.36, with an average non-optimum factor of 1.68. Application of the refined weights model yields non-optimum factors between 1.00 and 1.45, with an average non-optimum factor of 1.13. To demonstrate their utility, these calculated non-optimum factors are then used to predict more realistic (i.e., heavier) gore weights for a proposed heavy-lift launch vehicle's propellant tank bulkheads.
Low Cost Manufacturing of Composite Cryotanks
This viewgraph presentation reviews research and development of cryotank manufacturing conducted by Northrup Grumman. The objectives of the research and development included the development and validation of manufacturing processes and technology for fabrication of large scale cryogenic tanks, the establishment of a scale-up and facilitization plan for full scale cryotanks, the development of non-autoclave composite manufacturing processes, the fabrication of subscale tank joints for element tests, the performance of manufacturing risk reduction trials for the subscale tank, and the development of full-scale tank manufacturing concepts.