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The mechanical performance of the over-the-wing engine is described with emphasis on the advanced technology components. The overall dynamic response of the engine was excellent.
Compliant thermal barriers play a pivotal role in the thermal protection systems of advanced aerospace vehicles. Both the thermal properties and mechanical performance of these barriers are critical in determining their successful implementation. Due to the custom nature of many thermal barriers, designers of advanced spacecraft have little guidance as to the design, selection, and implementation of these elements. As part of an effort to develop a more fundamental understanding of the interrelationship between thermal barrier design and performance, mechanical testing of thermal barriers was conducted. Two different types of thermal barriers with several core insulation density levels ranging from 62 to 141 kg/cu m were investigated. Room-temperature compression tests were conducted on samples to determine load performance and assess thermal barrier resiliency. Results showed that the loading behavior of these thermal barriers was similar to other porous, low-density, compliant materials, such as elastomeric foams. Additionally, the insulation density level had a significant non-linear impact on the stiffness and peak loads of the thermal barriers. In contrast, neither the thermal barrier type nor the level of insulation density significantly influenced the room-temperature resiliency of the samples.
The National Aeronautics and Space Administration has been developing a novel docking system to meet the requirements of future exploration missions to low-Earth orbit and beyond. A dynamic gas pressure seal is located at the main interface between the active and passive mating components of the new docking system. This seal is designed to operate in the harsh space environment, but is also to perform within strict loading requirements while maintaining an acceptable level of leak rate. In this study, a candidate silicone elastomer seal was designed, and multiple subscale test articles were manufactured for evaluation purposes. The force required to fully compress each test article at room temperature was quantified and found to be below the maximum allowable load for the docking system. However, a significant amount of scatter was observed in the test results. Due to the stochastic nature of the mechanical performance of this candidate docking seal, a statistical process control technique was implemented to isolate unusual compression behavior from typical mechanical performance. The results of this statistical analysis indicated a lack of process control, suggesting a variation in the manufacturing phase of the process. Further investigation revealed that changes in the manufacturing molding process had occurred which may have influenced the mechanical performance of the seal. This knowledge improves the chance of this and future space seals to satisfy or exceed design specifications.
Carbon-carbon composite materials offer greater thermal efficiency, stiffness to weight ratio, tailorability, and dimensional stability than aluminum. These lightweight thermal materials could significantly reduce the overall costs associated with satellite thermal control and weight. However, the high cost and long lead-time for carbon-carbon manufacture have limited their widespread usage. Consequently, an informal partnership between government and industrial personnel called the Carbon-Carbon Spacecraft Radiator Partnership (CSRP) was created to foster carbon-carbon composite use for thermally and structurally demanding space radiator applications. The first CSRP flight opportunity is on the New Millennium Program (NMP) Earth Orbiter-1 (EO-1) spacecraft, scheduled for launch in late 1999. For EO-1, the CSRP designed and fabricated a Carbon-Carbon Radiator (CCR) with carbon-carbon facesheets and aluminum honeycomb core, which will also serve as a structural shear panel. While carbon-carbon is an ideal thermal candidate for spacecraft radiators, in practice there are technical challenges that may compromise performance. In this work, the thermal and mechanical performance of the EO-1 CCR is assessed by analysis and testing. Both then-nal and mechanical analyses were conducted to predict the radiator response to anticipated launch and on-orbit loads. The thermal model developed was based on thermal balance test conditions. The thermal analysis was performed using SINDA version 4.0. Structural finite element modeling and analysis were performed using SDRC/1-DEAS and UAI/NASTRAN, respectively. In addition, the CCR was subjected to flight qualification thermal/vacuum and vibration tests. The panel meets or exceeds the requirements for space flight and demonstrates promise for future satellite missions.
Nb-1Zr is a readily fabricable, dispersion-strengthened niobium alloy with improved elevated temperature properties compared to commercially pure niobium. This low to moderate strength niobium alloy is desirable in thermal management systems for fission power and propulsion applications; however, pre-alloyed spherical feedstock for powder-based additive manufacturing processes is generally unavailable. Laser powder bed fusion (L-PBF) in-situ alloying of a Nb-1Zr chemistry was performed using a high-purity niobium feedstock coated with 1 wt% of zirconium-based ceramic compound nano-powders: ZrO 2 , ZrC, and ZrH 2 . The additively manufactured material was mechanically tested at elevated temperatures within a controlled environment in the as-built condition. For comparison, traditionally manufactured wrought feedstock was also tested within an identical environment. Examination of the materials’ microstructure and short-term elevated temperature mechanical response determined: in-situ alloying of spherical Nb powder feedstock with Zr-based ceramic nanoparticles can be successfully implemented; the ZrO 2 addition offered the greatest elevated temperature stability and strength but exhibited brittle behavior; the ZrC addition achieved a balance of room temperature ductility and elevated temperature strength; and the ZrH 2 -modified material had the lowest as-built interstitial content and showed high ductility with mechanical performance most similar to wrought Nb-1Zr.
Results of initial tests of the under the wing experimental engine and boilerplate nacelle are presented. The mechanical performance of the engine is reported with emphasis on the advanced technology components. Technology elements of the propulsion system covered include: system dynamics, composite fan blades, reduction gear, lube and accessory drive system, fan frame, inlet, core cowl cooling, fan exhaust nozzle, and digital control system.
The utilization of lattice-type cellular architectures has seen a significant increase, owing to their predictable shape and the ability to fabricate templated porous materials through low-cost 3D-printing methods. Frames based on atomic lattice structures such as face-centered cubic (FCC), body-centered cubic (BCC), or simple cubic (SC) have been utilized. In FDM, the mechanical performance has been impeded by stress concentration at the nodes and melt-solidification interfaces arising from layer-by-layer deposition. Adding plates to the frames has resulted in improvements with a concurrent increase in weight and hot-pocket-induced dimensional impact in the closed cells formed. In this paper, we explore compressive performance from the partial addition of plates to the frames of a SC-BCC lattice. Compression testing of both single unit cells and 4 × 4 × 4 lattices in all three axial directions is conducted to examine stress transfer to the nearest neighbor and assess scale-up stress transfer. Our findings reveal that hybrid lattice structure unit cells exhibit significantly improved modulus in the range of 125% to 393%, specific modulus in the range of 13% to 120%, and energy absorption in the range of 17% to 395% over the open lattice. The scaled-up lattice modulus increased by 8% to 400%, specific modulus by 2% to 107%, and energy absorption by 37% to 553% over the lattice frame. Parameters that emerged as key to improved lightweighting.
A lightweight sandwich support structure, for the combustor chamber of a new generation liquid propellant rocket engine, was designed and fabricated using a polymer matrix composite (PMC) facesheet on a Ti honeycomb core. The PMC facesheet consisted of high stiffness carbon fiber, M40JB, and high temperature Polyimides, such as PMR-II-50 and HFPE-II-52. Six different fiber architectures; four harness satin (4HS) woven fabric, uni-tape, woven-uni hybrid, stitched woven fabric, stitched uni-tape and triaxial braided structures have been investigated for optimum stiffness-thickness-weight-hygrothermal performance design criteria for the hygrothermal-mechanical propulsion service exposure conditions including rapid heating up to 200 F/sec, maximum operating temperature of 600 F, internal pressure up to 100 psi. One of the specific objectives in this study is to improve composite blistering resistance in z-direction at minimum expense of in-plane mechanical properties. An extensive property-performance database including dry-wet mechanical properties at various temperatures, thermal-physical properties, such as blistering onset condition was generated for fiber architecture down-selection and design guidelines. Various optimized process methods such as vacuum bag compression molding, solvent assistant resin transfer molding (SaRTM), resin film infusion (RFI) and autoclaving were utilized for PMC panel fabrication depending on the architecture type. In the case of stitched woven fabric architecture, the stitch pattern in terms of stitch density and yarn size was optimized based on both in-plane mechanical properties and blistering performance. Potential reduction of the in-plane properties transverse to the line of stitching was also evaluated. Efforts have been made to correlate the experimental results with theoretical micro-mechanics predictions. Changes in deformation mechanism and failure sequences in terms of fiber architecture will be discussed.
Mechanical Testing of an advanced thermoplastic polyimide (LaRC-TM-SI) with known variations in molecular weight was performed over a range of temperatures below the glass transition temperature. The physical characterization, elastic properties and notched tensile strength were all determined as a function of molecular weight and test temperature. It was shown that notched tensile strength is a strong function of both temperature and molecular weight, whereas stiffness is only a strong function of temperature. A critical molecular weight (Mc) was observed to occur at a weight-average molecular weight (Mw) of approx. 22000 g/mol below which, the notched tensile strength decreases rapidly. This critical molecular weight transition is temperature-independent. Furthermore, inelastic analysis showed that low molecular weight materials tended to fail in a brittle manner, whereas high molecular weight materials exhibited ductile failure. The microstructural images supported these findings.
A rigid lightweight sandwich support structure, for the combustor chamber of a new generation liquid propellant rocket engine, was designed and fabricated using Polymer Matric Composite (PMC) facesheet on a Ti honeycomb or Carbon foam core. The facesheet consisted of high stiffness carbon fiber, M40JB, and high temperature Polyimides, such as PMR-11-50 and HFPE-II. Six different fiber architectures; 4HS woven fabric, uni-fabric, woven-uni hybrid, stitched woven fabric, stitched uni-fabric and tri-axial braided structures have been investigated for optimum stiffness-thickness-weight-performance design criteria for the hygrothermal-mechanical propulsion service exposure conditions including rapid heating up to 200 F/sec, maximum operating temperature of 600 F, internal pressure up to 100 psi. An extensive property and performance database including dry-wet mechanical properties at both 25 F and 600 F in various loading modes, thermal and physical properties including blistering onset condition was developed for fiber architecture down-selection and design guidelines. Various optimized process methods including vacuum bag compression molding, solvent assistant RTM (SaRTM), resin film infusion (RFI) were utilized for PMC panel fabrication depending on the architecture type. In the case of stitched woven fabric architecture, the optimal stitch pattern was chosen in terms of stitch density and yarn size, based on both in-plane mechanical properties and blistering performance. Potential reduction of the in-plane properties transverse to the line of stitching was also evaluated. Attempt to correlate the experimental results with theoretical micro-mechanics predictions will be presented.
The National Aeronautics and Space Administration is developing a Low Impact Docking System (LIDS) for future exploration missions. The mechanism is a new state-of-the-art device for in-space assembly of structures and rendezvous of vehicles. At the interface between two pressurized modules, each with a version of the LIDS attached, a composite elastomer-metal seal assembly prevents the breathable air from escaping into the vacuum of space. Attached to the active LIDS, this seal mates against the passive LIDS during docking operation. The main interface seal assembly must exhibit low leak and outgas values, must be able to withstand various harsh space environments, must remain operational over a range of temperatures from -50 C to 75 C, and perform after numerous docking cycles. This paper presents results from a comprehensive study of the mechanical performance of four candidate subscale seal assembly designs at -50, 23, 50, and 75 C test temperatures. In particular, the force required to fully compress the seal during docking, and that which is required for separation during the undocking operation were measured. The height of subscale main interface seal bulbs, as well as the test temperature, were shown to have a significant effect on the forces the main interface seal of the LIDS may experience during docking and undocking operations. The average force values required to fully compress each of the seal assemblies were shown to increase with test temperature by approximately 50% from -50 to 75 C. Also, the required compression forces were shown to increase as the height of the seal bulb was increased. The seal design with the tallest elastomer seal bulb, which was 31% taller than that with the shortest bulb, required force values approximately 45% higher than those for the shortest bulb, independent of the test temperature. The force required to separate the seal was shown to increase with decreasing temperature after 15 hours of simulated docking. No adhesion force was observed at 75 C, while magnitudes of up to 235 lbf were recorded at the refrigerated temperature. In addition, the adhesion force was observed to increase with bulb height. When compared with the LIDS program requirements, the measured compression force values were found to be below the maximum allowable load allotted to the main interface seal. However, the measured adhesion force values at the refrigerated test temperature were found to exceed the program limits.
The performance test results of the final under-the-wing engine configuration are presented. One hundred and six hours of engine operation were completed, including mechanical and performance checkout, baseline acoustic testing with a bellmouth inlet, reverse thrust testing, acoustic technology tests, and limited controls testing. The engine includes a variable pitch fan having advanced composite fan blades and using a ball-spline pitch actuation system.
The microstructure and mechanical properties of additively manufactured (AM) alloys can be significantly affected by variations in cooling rates, resulting from different process conditions across different additive manufacturing (AM) platforms. Therefore, it is crucial to understand the effect of manufacturing process on the microstructure and mechanical properties of AM Inconel 718. This study examines three AM processes: laser powder bed fusion, laser powder directed energy deposition, and wire arc additive manufacturing. Results show that fully heat treated laser powder bed fused (L-PBF) and wire arc additively manufactured (WAAM) Inconel 718 specimens exhibit higher strength compared to laser powder directed energy deposited (LP-DED) ones due to finer grain structure in L-PBF and retained dendritic microstructure in WAAM. The ductility in LP-DED Inconel 718 was slightly higher compared to WAAM and L-PBF due to relatively small carbide size, which causes stress concentration in a small material volume, leading to delayed fracture.
Nb-1Zr is a readily-fabricable, dispersion-strengthened niobium alloy with improved elevated temperature properties compared to commercially pure niobium. This moderate strength niobium alloy is desirable in thermal management systems for fission power and propulsion applications; however, pre-alloyed spherical feedstock is generally unavailable. Laser powder bed fusion (L-PBF) in-situ alloying of a Nb-1Zr chemistry was performed using a pure niobium feedstock coated with 1 wt% of zirconium-based ceramic compound nano-powders. The additively manufactured material was mechanically tested at elevated temperatures in a controlled environment in the as-built condition. For comparison, traditionally manufactured wrought feedstock was also tested under identical conditions. Examination of the materials’ microstructure and elevated temperature mechanical response aimed to determine the Zr-compound nano-powder addition with the best performance, and the overall feasibility of the in-situ alloying approach.
Three-dimensional (3D) woven preforms are currently being considered for use as primary structural components. Lack of technology to properly manufacture, characterize and predict mechanical properties, and predict damage mechanisms leading to failure are problems facing designers of textile composite materials. Two material systems with identical specifications but different manufacturing approaches are investigated. One manufacturing approach resulted in an irregular (nonuniform) preform geometry. The other approach yielded the expected preform geometry (uniform). The objectives are to compare the mechanical properties of the uniform and nonuniform angle interlock 3D weave constructions. The effect of adding layers of laminated tape to the outer surfaces of the textile preform is also examined. Damage mechanisms are investigated and test methods are evaluated.
In the process of the present invention, a non-polar, aprotic solvent is removed from an oligomer/polymer solution by freeze-drying in order to produce IPNs and semi-IPNs. By thermally quenching the solution to a solid in a short length of time, the size of the minor constituent-rich regions is greatly reduced as they are excluded along with the major constituent from the regions of crystallizing solvent. The use of this process sequence of controlling phase morphology provides IPNs and semi-IPNs with improved fracture toughness, microcracking resistance, and other physical-mechanical properties as compared to IPNs and semi-IPNs formed when the solvent is evaporated rather than sublimed.
In the process of the present invention, a non-polar, aprotic solvent is removed from an oligomer/polymer solution by freeze-drying in order to produce IPNs and semi-IPNs. By thermally quenching the solution to a solid in a short length of time, the size of the minor constituent-rich regions is greatly reduced as they are excluded along with the major constituent from the regions of crystallizing solvent. The use of this process sequence of controlling phase morphology provides IPNs and semi-IPNs with improved fracture toughness, microcracking resistance, and other physical-mechanical properties as compared to IPNs and semi-IPNs formed when the solvent is evaporated rather than sublimed.
In response to the elevated-temperature and weight-reduction demands of modern aerospace applications, a novel oxide-dispersion-strengthened low-density niobium alloy (LDNb-ODS) was fabricated using laser powder bed fusion (L-PBF). To overcome powder procurement barriers, L-PBF feedstock was produced by blending commercial Nb521, Ti64, and Cr powder with Y2O3 nanoparticles via resonant acoustic mixing. Following L-PBF and a 1400°C vacuum heat treatment, the alloy achieved a density of 6.73 g/cc and a fine mean grain size of 4.62 µm stabilized by uniform ~30 nm yttria dispersoids. Microstructural analysis revealed a chemically inhomogeneous build with lack-of-fusion defects and a titanium (Ti) shift from a nominal 31.5 wt% in the starting powder blend to 24.8 wt% in the printed part due to preferential Ti loss during printing. Elevated-temperature tensile testing demonstrated that LDNb-ODS maintained a superior specific yield strength of 60-85 MPa/(g/cc) up to 800°C, outperforming nickel-based alloys Ni625, Ni230, and GRX-810. Between 870°C and 950°C, its specific strength surpassed both Ni718 and Ni625. In rapid stress-rupture testing at 1093°C and 20.7 MPa, uncoated LDNb-ODS survived 21.4 hours (a tenfold increase over legacy C-103) while an R512E silicide coating extended rupture life to 84.8 hours, confirming that oxidation accelerates low-stress failure. These findings demonstrate that additive manufacturing of LDNb-ODS provides a viable, lightweight alternative to nickel-based superalloys for high-temperature (>850°C) aerospace components.