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At least 991 records · Page 55

All sources NASA NTRS DOE OSTI

NASA NTRS1998

Study Methods to Standardize Thermography NDE

The purpose of this work is to develop thermographic inspection methods and standards for use in evaluating structural composites and aerospace hardware. Qualification techniques and calibration methods are investigated to standardize the thermographic method for use in the field. Along with the inspections of test standards structural hardware, support hardware is designed and fabricated to aid in the thermographic process. Also, a standard operating procedure is developed for performing inspections with the Bales Thermal Image Processor (TIP). Inspections are performed on a broad range of structural composites. These materials include various graphite/epoxies, graphite/cyanide-ester, graphite/silicon-carbide, graphite phenolic and Keviar/epoxy. Also metal honeycomb (titanium and aluminum faceplates over an aluminum honeycomb core) structures are investigated. Various structural shapes are investigated and the thickness of the structures vary from as few as 3 plies to as many as 80 plies. Special emphasis is placed on characterizing defects in attachment holes and bondlines, in addition to those resulting from impact damage and the inclusion of foreign matter. Image processing through statistical analysis and digital filtering is investigated to enhance the quality and quantify the NDE thermal images when necessary.

Walker, James L.↗

NASA NTRS1997

The Diagnostics of the External Plasma for the Plasma Rocket

Three regions of plasma temperature/energy are being investigated to understand fully the behavior of the plasma created by the propulsion device and the operation of the RPA. Each type of plasma has a RPA associated with it; i.e. a thermal RPA, a collimated RPA, and a high temperature RPA. Through the process of developing the thermal and collimated RPAs, the proper knowledge and experience has been gained to not only design a high temperature RPA for the plasma rocket, but to understand its operation, results, and uncertainty. After completing a literature search for, reading published papers on, and discussing the operation of the RPA with electric propulsion researchers, I applied the knowledge gained to the development of a RPA for thermal plasma. A design of a thermal RPA was made which compensates for a large Debye length and low ionized plasma. From this design a thermal RPA was constructed. It consists of an outer stainless steel casing, a phenolic insulator (outgases slightly), and stainless steel mesh for the voltage screens. From the experience and knowledge gained in the development of the thermal RPA, a RPA for collimated plasma was developed. A collimated RPA has been designed and constructed. It compensate for a smaller Debye length and much higher ionization than that existing in the thermal plasma. It is 17% of the size of the thermal RPA. A stainless steel casing shields the detector from impinging electrons and ions. An insulating material, epoxy resin, was utilized which has a negligible outgassing. This material can be molded in styrofoam and machined quite nicely. It is capable of withstanding moderately high temperatures. Attached to this resin insulator are inconel screens attached by silver plated copper wire to a voltage supply. All the work on the RPAs and thermal ion source, I performed in the University of Alabama in Huntsville's (UAH) engineering machine shop.

Karr, Gerald R.↗

NASA NTRS1998

Feasibility Assessment of Thermal Barrier Seals for Extreme Transient Temperatures

The assembly joints of modem solid rocket motor cases are generally sealed using conventional O-ring type seals. The 5500+ F combustion gases produced by rocket motors are kept a safe distance away from the seals by thick layers of phenolic insulation. Special compounds are used to fill insulation gaps leading up to the seals to prevent a direct flowpath to them. Design criteria require that the seals should not experience torching or charring during operation, or their sealing ability would be compromised. On limited occasions, NASA has observed charring of the primary O-rings of the Space Shuttle solid rocket nozzle assembly joints due to parasitic leakage paths opening up in the gap-fill compounds during rocket operation. NASA is investigating different approaches for preventing torching or charring of the primary O-rings. One approach is to implement a braided rope seal upstream of the primary O-ring to serve as a thermal barrier that prevents the hot gases from impinging on the O-ring seals. This paper presents flow, resiliency, and thermal resistance for several types of NASA rope seals braided out of carbon fibers. Burn tests were performed to determine the time to burn through each of the seals when exposed to the flame of an oxyacetylene torch (5500 F), representative of the 5500 F solid rocket motor combustion temperatures. Rope seals braided out of carbon fibers endured the flame for over six minutes, three times longer than solid rocket motor burn time. Room and high temperature flow tests are presented for the carbon seals for different amounts of linear compression. Room temperature compression tests were performed to assess seal resiliency and unit preloads as a function of compression. The thermal barrier seal was tested in a subscale "char" motor test in which the seal sealed an intentional defect in the gap insulation. Temperature measurements indicated that the seal blocked 2500 F combustion gases on the upstream side with very little temperature rise on the downstream side.

Steinetz, Bruce M.↗

NASA NTRS1998

Study Methods to Standardize Thermography NDE

The purpose of this work is to develop thermographic inspection methods and standards for use in evaluating structural composites and aerospace hardware. Qualification techniques and calibration methods are investigated to standardize the thermographic method for use in the field. Along with the inspections of test standards structural hardware, support hardware is designed and fabricated to aid in the thermographic process. Also, a standard operating procedure is developed for performing inspections with the Bales Thermal Image Processor (TIP). Inspections are performed on a broad range of structural composites. These materials include graphite/epoxies, graphite/cyanide-ester, graphite/silicon-carbide, graphite phenolic and Kevlar/epoxy. Also metal honeycomb (titanium and aluminum faceplates over an aluminum honeycomb core) structures are investigated. Various structural shapes are investigated and the thickness of the structures vary from as few as 3 plies to as many as 80 plies. Special emphasis is placed on characterizing defects in attachment holes and bondlines, in addition to those resulting from impact damage and the inclusion of foreign matter. Image processing through statistical analysis and digital filtering is investigated to enhance the quality and quantify the NDE thermal images when necessary.

Walker, James L.↗

NASA NTRS1975

Experimental Evaluation of Silica as a Thermal Protection System for Planetary Entry Probes

The evolution and pursuit of the reflecting heat-shield concept has resulted in a refined and sophisticated high-performance material: ultra-pure, slip-cast, fused silica (SCFS). Silica is "tailored" to the intense planetary entry environments through a combination of its good ablative qualities and outstanding reflective properties in the correct spectral region. Extensive development of SCFS has produced a series of candidate materials which are experimentally evaluated and compared in a simulated entry environment. The detailed response of silica to combined heating is discussed, and the relative merits of SCFS and carbon phenolic are indicated.

David L Peterson↗

NASA NTRS1999

Development of a Pressure-Fed Rocket Engine Using Hydrogen Peroxide and JP-8

A pressure-fed rocket engine using hydrogen peroxide and JP-8 was designed and demonstrated for upper stage space transportation applications. The engine utilizes silver plated screen catalyst to decompose 85% hydrogen peroxide (by weight). The decomposed high-temperature gas causes the JP-8 to auto-ignite and undergo chemical reaction. An ablative chamber using silica phenolic was used as the combustion chamber. The present effort includes catalyst development, injector evaluation, and chamber char and erosion characterization. Wagonwheel-type distribution and support plates were used in the catalyst bed design to reduce pressure drop, to increase active area, and to avoid flow channeling. Various fuel injection and mixing approaches were investigated and the combustion efficiency was in the range from 89 to 98%, depending on injector design and operating conditions. Fuel film cooling, up to 40% of the total fuel flow, was found to be ineffective with the current engine configuration. A ring injector was selected because of the satisfactory performance and its low cost. The char and erosion rates were determined through hot-fire tests. The char depth was found to be a function of axial distance, which was related to the progress of chemical reaction. Correlation of the char and erosion rates was developed as a design database for future applications,

Wu, P.-K.↗

NASA NTRS1999

Alternatives to Silicon Rubber Thermal Barrier in RSRM Nozzle Joints

Thiokol Propulsion encountered a situation where gas path can form in the Filler material in joints between the phenolic in the Space Shuttle Reusable Solid Rocket Motor nozzle. An investigation determined that the method for injecting the filler material into the joint created a pressurized pocket of air that forced its way to the surface while the filler was curing. The path that the air took remained open some situations leaving a path for hot gas to get deep into the joint. The condition that created the gas path has been corrected by modifications to the manufacturing process, but a design solution that would reduce processing time is being pursued. A strip of open cell installed deep into the joint, can serve as a barrier to the Filler material. Testing indicates that the foam eliminates the pressure pocket, and does not prevent detection of leakage of the o-ring seals during leak test. Open cell foam does not have thermal resistance capability and mechanical structure, which would be required in the event of a gas path. A replacement to open cell foam is now being considered. Braided carbon fiber rope has the permeability characteristics of foam and it has thermal resistance characteristics as well. Braided carbon fiber rope (BCFR) has been tested in simulated joints in subscale rocket motor. Results of the testing indicate that braided carbon fiber rope can significantly reduce gas temperature by as much es 3000 F in a nozzle joint configuration. BCFR can be installed as a barrier to stop the joint filler material or it can used as the thermal barrier alone.

Totman, Peter D.↗

NASA NTRS1998

Tooling Foam for Structural Composite Applications

Tooling technology applications for composite structures fabrication have been expanded at MSFC's Productivity Enhancement Complex (PEC). Engineers from NASA/MSFC and Lockheed Martin Corporation have developed a tooling foam for use in composite materials processing and manufacturing that exhibits superior thermal and mechanical properties in comparison with other tooling foam materials. This tooling foam is also compatible with most preimpregnated composite resins such as epoxy, bismaleimide, phenolic and their associated cure cycles. MARCORE tooling foam has excellent processability for applications requiring either integral or removable tooling. It can also be tailored to meet the requirements for composite processing of parts with unlimited cross sectional area. A shelf life of at least six months is easily maintained when components are stored between 50F - 70F. The MARCORE tooling foam system is a two component urethane-modified polyisocyanurate, high density rigid foam with zero ozone depletion potential. This readily machineable, lightweight tooling foam is ideal for composite structures fabrication and is dimensionally stable at temperatures up to 350F and pressures of 100 psi.

DeLay, Tom↗

NASA NTRS1998

RSRM Nozzle Anomalous Throat Erosion Investigation Overview

In September, 1996, anomalous pocketing erosion was observed in the aft end of the throat ring of the nozzle of one of the reusable solid rocket motors (RSRM 56B) used on NASA's space transportation system (STS) mission 79. The RSRM throat ring is constructed of bias tape-wrapped carbon cloth/ phenolic (CCP) ablative material. A comprehensive investigation revealed necessary and sufficient conditions for occurrence of the pocketing event and provided rationale that the solid rocket motors for the subsequent mission, STS-80, were safe to fly. The nozzles of both of these motors also exhibited anomalous erosion similar to, but less extensive than that observed on STS-79. Subsequent to this flight, the investigation to identify both the specific causes and the corrective actions for elimination of the necessary and sufficient conditions for the pocketing erosion was intensified. A detailed fault tree approach was utilized to examine potential material and process contributors to the anomalous performance. The investigation involved extensive constituent and component material property testing, pedigree assessments, supplier audits, process audits, full scale processing test article fabrication and evaluation, thermal and thermostructural analyses, nondestructive evaluation, and material performance tests conducted using hot fire simulation in laboratory test beds and subscale and full scale solid rocket motor static test firings. This presentation will provide an over-view of the observed anomalous nozzle erosion and the comprehensive, fault-tree based investigation conducted to resolve this issue.

Clinton, R. G., Jr.↗

NASA NTRS1999

Structural Evaluation of the RSRM Nozzle Replacement Adhesive

This paper describes the structural performance evaluation of a replacement adhesive for the Reusable Solid Rocket Motor (RSRM) nozzle utilizing finite element analysis. Due to material obsolescence and industrial safety issues, the two current structural adhesives, EA 913 and EA 946 are to be replaced with a new adhesive. TIGA 321. The structural evaluation in support of the adhesive replacement effort includes residual stress, transportation, and flight analyses. Factors of safety are calculated using the stress response from each analysis. The factors of safety are used as the limiting criteria to compare the replacement adhesive against the current adhesives. Included in this paper are the analytical approach, assumptions and modeling techniques as well as the results of the evaluation. An important factor to the evaluation is the similarity in constitutive material properties (elastic modulus and Poisson's ratio) between TIGA 321 and EA 913. This similarity leads to equivalent material response from the two adhesives. However, TIGA 321 surpasses EA 913's performance due to higher material capabilities. Conversely, the change in stress response from EA 946 to TIGA 321 is more apparent: this is primarily attributed to the difference in the modulii of the two adhesives, which differ by two orders of magnitude. The results of the bondline evaluation indicate that the replacement adhesive provides superior performance than the current adhesives with only minor exceptions. Furthermore, TIGA 321 causes only a minor chance in the response of the phenolic and metal components.

Batista-Rodriguez, A.↗

NASA NTRS1999

Feasibility Assessment of Thermal Barriers for RSRM Nozzle Joint Locations

Solid rockets, including the Space Shuttle solid rocket motor, are generally manufactured in large segments which are then shipped to their final destination where they are assembled. These large segments are sealed with a system of primary and secondary 0-rings to contain combustion gases inside the rocket which are at pressures of up to 900 psi and temperatures of up to 5500 F. The seals are protected from hot combustion gases by thick layers of phenolic insulation and by joint-filling compounds between these layers. Recently, though, routine inspections of nozzle-to-case joints in the Shuttle solid rocket motors during disassembly revealed erosion of the primary O-rings. Jets of hot gas leaked through gaps in the joint-filling compound between the layers of insulation and impinged on the O-rings. This is not supposed to take place, so NASA and Thiokol, the manufacturer of the rockets, initiated an investigation and found that design improvements could be made in this joint. One such improvement would involve using NASA Lewis braided thermal barriers as another level of protection for the O-ring seals against the hot combustion gases.

Steinetz, Bruce M.↗

NASA NTRS1999

Thermal Characterization of Adhesive

The current Space Shuttle Reusable Solid Rocket Motor (RSRM) nozzle adhesive bond system is being replaced due to obsolescence. Down-selection and performance testing of the structural adhesives resulted in the selection of two candidate replacement adhesives, Resin Technology Group's Tiga 321 and 3M's EC2615XLW. This paper describes rocket motor testing of these two adhesives. Four forty-pound charge motors were fabricated in configurations that would allow side by side comparison testing of the candidate replacement adhesives and the current RSRM adhesives. The motors provided an environment where the thermal performance of adhesives in flame surface bondlines was compared. Results of the FPC testing show that: 1) The phenolic char depths on radial bond lines is approximately the same and vary depending on the position in the blast tube regardless of which adhesive was used; 2) The adhesive char depth of the candidate replacement adhesives is less than the char depth of the current adhesives; 3) The heat-affected depth of the candidate replacement adhesives is less than the heat-affected depth of the current adhesives; and 4) The ablation rates for both replacement adhesives are slower than that of the current adhesives.

Spomer, Ken A.↗

NASA NTRS2000

NARC Rayon Replacement Program for the Space Shuttle Reusable Solid Rocket Motor Nozzle: Screening Summary

Thiokol Corporation and NASA MSFC are jointly developing a replacement for North American Rayon Corporation (NARC) Aerospace Grade Rayon (1650/720 continuous filament), the precursor for the Carbon Cloth Phenolic (CCP) ablatives used in the Space Shuttle Reusable Solid Rocket Motor (RSRM) Nozzles. NARC discontinued production of Aerospace Grade Rayon in September 1997. NASA maintains a stockpile of NARC Rayon to support RSRM production through the summer of 2005. The program plan for selection and qualification of a replacement for NARC rayon was approved in August 1998. Screening activities began in February 1999. The intent of this paper is to provide a summary of the data generated during the screening phase of the NARC Rayon Replacement Program. Twelve cellulose based fibers (rayon and lyocell) were evaluated. These fibers were supplied by three independent vendors. Many of these fibers were carbonized by two independent carbonizers. Each candidate was tested according to standard acceptance test methods at each step of the manufacturing process. Additional testing was performed with the candidate CCPS, including hot fire tests, Process studies and mechanical and thermal characterization. Six of the twelve fiber candidates tested were dropped at the conclusion of Phase 1. The reasons for the elimination of these candidates included; difficulties in processing the material in the whitegoods, carbon and CCP forms; poor composite mechanical performance; and future availability concerns. The remaining six fibers demonstrated enough promise to merit continued evaluation and optimization of the CCP fabrication process. Note: Certain CCP data falls under the restrictions of US export laws, (ITAR, etc.) and will not be included in this paper.

Cook, R. V.↗

NASA NTRS2000

Upper Stage Flight Experiment 10K Engine Design and Test Results

A 10,000 lbf thrust chamber was developed for the Upper Stage Flight Experiment (USFE). This thrust chamber uses hydrogen peroxide/JP-8 oxidizer/fuel combination. The thrust chamber comprises an oxidizer dome and manifold, catalyst bed assembly, fuel injector, and chamber/nozzle assembly. Testing of the engine was done at NASA's Stennis Space Center (SSC) to verify its performance and life for future upper stage or Reusable Launch Vehicle applications. Various combinations of silver screen catalyst beds, fuel injectors, and combustion chambers were tested. Results of the tests showed high C* efficiencies (97% - 100%) and vacuum specific impulses of 275 - 298 seconds. With fuel film cooling, heating rates were low enough that the silica/quartz phenolic throat experienced minimal erosion. Mission derived requirements were met, along with a perfect safety record.

Ross, R.↗

NASA NTRS2000

Development of Thermal Barriers For Solid Rocket Motor Nozzle Joints

Joints in the Space Shuttle solid rocket motors are sealed by O-rings to contain combustion gases inside the rocket that reach pressures of up to 900 psi and temperatures of up to 5500 F. To provide protection for the O-rings, the motors are insulated with either phenolic or rubber insulation. Gaps in the joints leading up to the O-rings are filled with polysulfide joint-fill compounds as an additional level of protection. The current RSRM nozzle-to-case joint design incorporating primary, secondary, and wiper O-rings experiences gas paths through the joint-fill compound to the innermost wiper O-ring in about one out of every seven motors. Although this does not pose a safety hazard to the motor, it is an undesirable condition that NASA and rocket manufacturer Thiokol want to eliminate. Each nozzle-to-case joint gas path results in extensive reviews and evaluation before flights can be resumed. Thiokol and NASA Marshall are currently working to improve the nozzle-to-case joint design by implementing a more reliable J-leg design that has been used successfully in the field and igniter joint. They are also planning to incorporate the NASA Glenn braided carbon fiber thermal barrier into the joint. The thermal barrier would act as an additional level of protection for the O-rings and allow the elimination of the joint-fill compound from the joint.

Steinetz, Bruce M.↗

NASA NTRS1994

Aerothermodynamics of Pyrolizing Surfaces in Hypersonic Rarefied Flows

Direct simulation Monte Carlo (DSMC) calculations of rarefied flows about entry bodies typically employ a fixed surface temperature or a radiative-equilibrium energy balance to compute that temperature. Such boundary conditions neglect any effects of heat capacitance and heat conduction in the spacecraft heat shield and, therefore, provide an upper bound for the surface temperature. Such calculations also neglect pyrolysis from the heat shield which can be significant for a high-energy incident flow at very low densities. Accurate prediction of both heating and aerodynamic forces requires including pyrolysis and surface heat transfer in the models for gas-surface interaction employed in DSMC methods. Although these physical models have long appeared in various continuum flow calculation codes, they have only recently appeared in DSMC codes which are required to simulate rarefied flows during entry at high altitudes. In the current implementation, routines from the widely distributed Charring Material Thermal Response and Ablation (CMA) program are coupled into a DSMC code to calculate the one-dimensional heat transfer into the carbon phenolic heat shield at each point on a vehicle surface. Temperature-dependent material properties, surface re-radiation, and in-depth pyrolysis were included in the calculation, but surface ablation was neglected. Sample calculations for entry of the Galileo probe into the atmosphere of Jupiter demonstrate that including pyrolysis in the model leads to significant differences in predicted aerodynamics. Granted, the drag coefficient does not depend strongly on the surface temperature which can itself be significantly below the radiative equilibrium value during entry. However, the surface mass flux due to pyrolysis of the material is significant once the probe drops to altitudes characterized by transition flow. This leads to a noticeable increase in drag and a decrease in heating compared to a body without pyrolysis.

Haas, Brian L.↗

NASA NTRS2001

Method of Preparing Polymers with Low Melt Viscosity

This invention is an improvement in standard polymerizations procedures, i.e., addition-type and step-growth type polymerizations, wherein monomers are reacted to form a growing polymer chain. The improvement includes employing an effective amount of a trifunctional monomer (such as a trifunctional amine anhydride, or phenol) in the polymerization procedure to form a mixture of polymeric materials consisting of branced polymers, star-shaped polymers, and linear polymers. This mixture of polymeric materials has a lower melt temperature and a lower melt viscosity than corresponding linear polymeric materials of equivalent molecular weight.

Jensen, Brian J.↗

NASA NTRS1994

Light Weight Ceramic Ablators for Mars Follow-on Mission Vehicle Thermal Protection System

New Light Weight Ceramic Ablators (LCA) were produced by using ceramic and carbon fibrous substrates, impregnated with silicone and phenolic resins. The special infiltration techniques (patent pending) were developed to control the amount of organic resins in the highly porous fiber matrices so that the final densities of LCA's range from 0.22 to 0.24 g/cc. This paper presents the thermal and ablative performance of the Silicone Impregnated Reusable Ceramic Ablators (SIRCA) in simulated entry conditions for Mars-Pathfinder in the Ames 60 MW Interaction Heating Facility (I HF). Arc jet test results yielded no evidence of char erosion and mass loss at high stagnation pressures to 0.25 atm. Minimal silica melt was detected on surface char at a stagnation pressure of 0.31 atm. Four ceramic substrates were used in the production of SIRCA's to obtain the effective of boron oxide present in substrate so the thermal performance of SIRCA's. A sample of SIRCA was also exposed to the same heating condition for five cycles and no significant mass loss or recession was observed. Tensile testing established that the SIRCA tensile strength is about a factor of two higher than that of the virgin substrates. Thermogravimetric Analysis (TGA) of the char in nitrogen and air showed no evidence of free carbon in the char. Scanning Electron Microscopy of the post test sample showed that the char surface consists of a fibrous structure that was sealed with a thin layer of silicon oxide melt.

Tran, Huy K.↗