Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “Staging”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Performance estimates for space shuttle vehicles using a hydrogen or a methane fueled turboramjet powered first stage

Two- and three-stage (second stage expendable) shuttle vehicles, both having a hydrogen-fueled, turboramjet-powered first stage, are compared with a two-stage, VTOHL, all-rocket shuttle in terms of payload fraction, inert weight, development cost, operating cost, and total cost. All of the vehicles place 22,680 kilograms of payload into a 500-kilometer orbit. The upper stage(s) uses hydrogen-oxygen rockets. The effect on payload fraction and vehicle inert weight of methane and methane-FLOX as a fuel-propellant combination for the three-stage vehicle is indicated. Compared with a rocket first stage for a two-stage shuttle, an airbreathing first stage results in a higher payload fraction and a lower operating cost, but a higher total cost. The effect on cost of program size and first-stage flyback is indicated. The addition of an expendable rocket second stage (three-stage vehicle) improves the payload fraction but is unattractive economically.

Knip, G., Jr.↗
NASA NTRS2007

Analytical Approach for Estimating Preliminary Mass of ARES I Crew Launch Vehicle Upper Stage Structural Components

In January 2004, President Bush gave the National Aeronautics and Space Administration (NASA) a vision for Space Exploration by setting our sight on a bold new path to go back to the Moon, then to Mars and beyond. In response to this vision, NASA started the Constellation Program, which is a new exploration launch vehicle program. The primary mission for the Constellation Program is to carry out a series of human expeditions ranging from Low Earth Orbit to the surface of Mars and beyond for the purposes of conducting human exploration of space, as specified by the Vision for Space Exploration (VSE). The intent is that the information and technology developed by this program will provide the foundation for broader exploration activities as our operational experience grows. The ARES I Crew Launch Vehicle (CLV) has been designated as the launch vehicle that will be developed as a "first step" to facilitate the aforementioned human expeditions. The CLV Project is broken into four major elements: First Stage, Upper Stage Engine, Upper Stage (US), and the Crew Exploration Vehicle (CEV). NASA's Marshall Space Flight Center (MSFC) is responsible for the design of the CLV and has the prime responsibility to design the upper stage of the vehicle. The US is the second propulsive stage of the CLV and provides CEV insertion into low Earth orbit (LEO) after separation from the First Stage of the Crew Launch Vehicle. The fully integrated Upper Stage is a mix of modified existing heritage hardware (J-2X Engine) and new development (primary structure, subsystems, and avionics). The Upper Stage assembly is a structurally stabilized cylindrical structure, which is powered by a single J-2X engine which is developed as a separate Element of the CLV. The primary structure includes the load bearing liquid hydrogen (LH2) and liquid oxygen (LOX) propellant tanks, a Forward Skirt, the Intertank structure, the Aft Skirt and the Thrust Structure. A Systems Tunnel, which carries fluid and electrical power functions to other Elements of the CLV, is included as secondary structure. The MSFC has an overall responsibility for the integrated US element as well as structural design an thermal control of the fuel tanks, intertank, interstage, avionics, main propulsion system, Reaction Control System (RCS) for both the Upper Stage and the First Stage. MSFC's Spacecraft and Vehicle Department, Structural and Analysis Design Division is developing a set of predicted mass of these elements. This paper details the methodology, criterion and tools used for the preliminary mass predictions of the upper stage structural assembly components. In general, weight of the cylindrical barrel sections are estimated using the commercial code Hypersizer, whereas, weight of the domes are developed using classical solutions. HyperSizer is software that performs automated structural analysis and sizing optimization based on aerospace methods for strength, stability, and stiffness. Analysis methods range from closed form, traditional hand calculations repeated every day in industry to more advanced panel buckling algorithms. Margin-of-safety reporting for every potential failure provides the engineer with a powerful insight into the structural problem. Optimization capabilities include finding minimum weight panel or beam concepts, material selections, cross sectional dimensions, thicknesses, and lay-ups from a library of 40 different stiffened and sandwich designs and a database of composite, metallic, honeycomb, and foam materials. Multiple different concepts (orthogrid, isogrid, and skin stiffener) were run for multiple loading combinations of ascent design load with and with out tank pressure as well as proof pressure condition. Subsequently, selected optimized concept obtained from Hypersizer runs was translated into a computer aid design (CAD) model to account for the wall thickness tolerance, weld land etc for developing the most probable weight of the components. The flow diram summarizes the analysis steps used in developing these predicted mass.

Aggarwal, Pravin↗
NASA NTRS1974

Single-stage experimental evaluation of tandem-airfoil rotor and stator blading for compressors, part 8

An experimental investigation was conducted with an 0.8 hub/tip ratio, single-stage, axial flow compressor to determine the potential of tandem-airfoil blading for improving the efficiency and stable operating range of compressor stages. The investigation included testing of a baseline stage with single-airfoil blading and two tandem-blade stages. The overall performance of the baseline stage and the tandem-blade stage with a 20-80% loading split was considerably below the design prediction. The other tandem-blade stage, which had a rotor with a 50-50% loading split, came within 4.5% of the design pressure rise (delta P(bar)/P(bar) sub 1) and matched the design stage efficiency. The baseline stage with single-airfoil blading, which was designed to account for the actual rotor inlet velocity profile and the effects of axial velocity ratio and secondary flow, achieved the design predicted performance. The corresponding tandem-blade stage (50-50% loading split in both blade rows) slightly exceeded the design pressure rise but was 1.5 percentage points low in efficiency. The tandem rotors tested during both phases demonstrated higher pressure rise and efficiency than the corresponding single-airfoil rotor, with identical inlet and exit airfoil angles.

Brent, J. A.↗
NASA NTRS1990

Advanced launch vehicle upper stages using liquid propulsion and metallized propellants

Metallized propellants are liquid propellants with a metal additive suspended in a gelled fuel or oxidizer. Typically, aluminum particles are the metal additives. These propellants provide increase in the density and/or the specific impulse of the propulsion system. Using metallized propellants for volume- and mass-constrained upper stages can deliver modest increases in performance for Low Earth Orbit to Geosynchronous Earth Orbit and other Earth orbital transfer missions. Metallized propellants, however, can enable very fast planetary missions with a single-stage upper stage system. Trade studies comparing metallized propellant stage performance with non-metallized upper stages and the Inertial Upper Stage are presented. These upper stages are both one- and two-stage vehicles that provide the added energy to send payloads to altitudes and onto trajectories that are unattainable with only the launch vehicle. The stage designs are controlled by the volume and the mass constraints of the Space Transportation System and Space Transportation System-Cargo launch vehicles. The influences of the density and specific impulse increases enabled by metallized propellants are examined for a variety of different stage and propellant combinations.

Palaszewski, B. A.↗
NASA NTRS1990

Advanced Launch Vehicle Upper Stages Using Liquid Propulsion and Metallized Propellants

Metallized propellants are liquid propellants with a metal additive suspended in a gelled fuel or oxidizer. Typically, aluminum (Al) particles are the metal additive. These propellants provide increase in the density and/or the specific impulse of the propulsion system. Using metallized propellant for volume-and mass-constrained upper stages can deliver modest increases in performance for low earth orbit to geosynchronous earth orbit (LEO-GEO) and other earth orbital transfer missions. Metallized propellants, however, can enable very fast planetary missions with a single-stage upper stage system. Trade studies comparing metallized propellant stage performance with non-metallized upper stages and the Inertial Upper Stage (IUS) are presented. These upper stages are both one- and two-stage vehicles that provide the added energy to send payloads to altitudes and onto trajectories that are unattainable with only the launch vehicle. The stage designs are controlled by the volume and the mass constraints of the Space Transportation System (STS) and Space Transportation System-Cargo (STS-C) launch vehicles. The influences of the density and specific impulse increases enabled by metallized propellants are examined for a variety of different stage and propellant combinations.

Palaszewski, Bryan A.↗
NASA NTRS1992

Upper stages using liquid propulsion and metallized propellants

Metallized propellants are liquid propellants with a metal additive suspended in a gelled fuel. Typically, aluminum particles are the metal additive. These propellants increase the density and/or the specific impulse of the propulsion system. Using metallized propellants for volume- and mass-constrained upper stages can deliver modest increases in performance for low Earth orbit to geosynchronous Earth orbit (LEO-GEO) and other Earth-orbital transfer missions. However, using metallized propellants for planetary missions can deliver great reductions in flight time with a single-stage, upper-stage system. Tradeoff studies comparing metallized propellant stage performance with nonmetallized upper stages and the Inertial Upper Stage (IUS) are presented. These upper stages, launched from the STS and STS-C, are both one- and two-stage vehicles that provide the added energy to send payloads to high altitude orbits and onto interplanetary trajectories that are unattainable with only the Space Transportation System (STS) and the Space Transportation System-Cargo (STS-C). The stage designs are controlled by the volume and the mass constraints of the STS and STS-C launch vehicles. The influences of the density and specific impulse increases enabled by metallized propellants are examined for a variety of different stage and propellant combinations.

Palaszewski, Bryan A.↗
NASA NTRS1991

Conceptual design of two-stage-to-orbit hybrid launch vehicle

The object of this design class was to design an earth-to orbit vehicle to replace the present NASA space shuttle. The major motivations for designing a new vehicle were to reduce the cost of putting payloads into orbit and to design a vehicle that could better service the space station with a faster turn-around time. Another factor considered in the design was that near-term technology was to be used. Materials, engines and other important technologies were to be realized in the next 10 to 15 years. The first concept put forth by NASA to meet these objectives was the National Aerospace Plane (NASP). The NASP is a single-stage earth-to-orbit air-breathing vehicle. This concept ran into problems with the air-breathing engine providing enough thrust in the upper atmosphere, among other things. The solution of this design class is a two-stage-to-orbit vehicle. The first stage is air-breathing and the second stage is rocket-powered, similar to the space shuttle. The second stage is mounted on the top of the first stage in a piggy-back style. The vehicle takes off horizontally using only air-breathing engines, flies to Mach six at 100,000 feet, and launches the second stage towards its orbital path. The first stage, or booster, will weigh approximately 800,000 pounds and the second stage, or orbiter will weigh approximately 300,000 pounds. The major advantage of this design is the full recoverability of the first stage compared with the present solid rocket booster that are only partially recoverable and used only a few times. This reduces the cost as well as providing a more reliable and more readily available design for servicing the space station. The booster can fly an orbiter up, turn around, land, refuel, and be ready to launch another orbiter in a matter of hours.

Source record↗
NASA NTRS2010

Two stage catalytic combustor

A catalytic combustor (14) includes a first catalytic stage (30), a second catalytic stage (40), and an oxidation completion stage (49). The first catalytic stage receives an oxidizer (e.g., 20) and a fuel (26) and discharges a partially oxidized fuel/oxidizer mixture (36). The second catalytic stage receives the partially oxidized fuel/oxidizer mixture and further oxidizes the mixture. The second catalytic stage may include a passageway (47) for conducting a bypass portion (46) of the mixture past a catalyst (e.g., 41) disposed therein. The second catalytic stage may have an outlet temperature elevated sufficiently to complete oxidation of the mixture without using a separate ignition source. The oxidation completion stage is disposed downstream of the second catalytic stage and may recombine the bypass portion with a catalyst exposed portion (48) of the mixture and complete oxidation of the mixture. The second catalytic stage may also include a reticulated foam support (50), a honeycomb support, a tube support or a plate support.

Alvin, Mary Anne↗
NASA NTRS2011

Ares I-X First Stage Separation Loads and Dynamics Reconstruction

The Ares I-X flight test provided NASA with the opportunity to test hardware and gather critical data to ensure the success of future Ares I flights. One of the primary test flight objectives was to evaluate the environment during First Stage separation to better understand the conditions that the J-2X second stage engine will experience at ignition [1]. A secondary objective was to evaluate the effectiveness of the stage separation motors. The Ares I-X flight test vehicle was successfully launched on October 29, 2009, achieving most of its primary and secondary test objectives. Ground based video camera recordings of the separation event appeared to show recontact of the First Stage and the Upper Stage Simulator followed by an unconventional tumbling of the Upper Stage Simulator. Closer inspection of the videos and flight test data showed that recontact did not occur. Also, the motion during staging was as predicted through CFD analysis performed during the Ares I-X development. This paper describes the efforts to reconstruct the vehicle dynamics and loads through the staging event by means of a time integrated simulation developed in TREETOPS, a multi-body dynamics software tool developed at NASA [2]. The simulation was built around vehicle mass and geometry properties at the time of staging and thrust profiles for the first stage solid rocket motor as well as for the booster deceleration motors and booster tumble motors. Aerodynamic forces were determined by models created from a combination of wind tunnel testing and CFD. The initial conditions such as position, velocity, and attitude were obtained from the Best Estimated Trajectory (BET), which is compiled from multiple ground based and vehicle mounted instruments. Dynamic loads were calculated by subtracting the inertial forces from the applied forces. The simulation results were compared to the Best Estimated Trajectory, accelerometer flight data, and to ground based video.

Demory, Lee↗
NASA NTRS2011

Design of a 3-Stage ADR for the Soft X-Ray Spectrometer Instrument on the Astro-H Mission

The Japanese Astro-H mission will include the Soft X-ray Spectrometer (SXS) instrument, whose 36-pixel detector array of ultra-sensitive x-ray microcalorimeters requires cooling to 50 mK. This will be accomplished using a 3-stage adiabatic demagnetization refrigerator (ADR). The design is dictated by the need to operate with full redundancy with both a superfluid helium dewar at 1.3 K or below, and with a 4.5 K Joule-Thomson (JT) cooler. The ADR is configured as a 2-stage unit that is located in a well in the helium tank, and a third stage that is mounted to the top of the helium tank. The third stage is directly connected through two heat switches to the JT cooler and the helium tank, and manages heat flow between the two. When liquid helium is present, the 2-stage ADR operates in a single-shot manner using the superfluid helium as a heat sink. The third stage may be used independently to reduce the time-average heat load on the liquid to extend its lifetime. When the liquid is depleted, the 2nd and 3rd stages operate as a continuous ADR to maintain the helium tank at as low a temperature as possible - expected to be 1.2 K - and the 1st stage cools from that temperature as a single-stage, single-shot ADR. The ADR s design and operating modes are discussed, along with test results of the prototype 3-stage ADR.

Shirron, Peter J.↗
NASA NTRS2015

Rho-Isp Revisited and Basic Stage Mass Estimating for Launch Vehicle Conceptual Sizing Studies

The ideal rocket equation is manipulated to demonstrate the essential link between propellant density and specific impulse as the two primary stage performance drivers for a launch vehicle. This is illustrated by examining volume-limited stages such as first stages and boosters. This proves to be a good approximation for first-order or Phase A vehicle design studies for solid rocket motors and for liquid stages, except when comparing to hydrogen-fueled stages. A next-order mass model is developed that is able to model the mass differences between hydrogen-fueled and other stages. Propellants considered range in density from liquid methane to inhibited red fuming nitric acid. Calculated comparisons are shown for solid rocket boosters, liquid first stages, liquid upper stages, and a balloon-deployed single-stage-to-orbit concept. The derived relationships are ripe for inclusion in a multi-stage design space exploration and optimization algorithm, as well as for single-parameter comparisons such as those shown herein.

Kibbey, Timothy P.↗
NASA NTRS1985

Small, two-stage, partial-admission turbine

The Rocketdyne Orbital Transfer Vehicle (OTV) cryogenic, rocket engine system, high pressure, liquid hydrogen turbopump was designed with a two-stage, partial-admission axial turbine. The turbine is basically two single-stage, partial-admission, subsonic impulse stages designed so the kinetic energy leaving the first-stage rotor is discharged directly into the second-stage nozzle at nominal operation to minimize staging losses. Very little data were available in the literature for this type of turbine design. Therefore, it was decided to test a full-size model of the turbine design using ambient-temperature gaseous nitrogen as the working fluid. The tester design features a variable orientation second-stage nozzle to determine the optimum circumferential location for highest performance. The tester also features the capability to vary the nozzle arcs of admission and incorporates quartz windows to study the flowfield upstream of the second-stage nozzle using a laser velocimeter. The test operations will probe the efficiency and flow characteristics for three arcs of admission and the effects of second-stage nozzle circumferential orientations over wide ranges of speed and pressure ratios as well as the interstage pressure distributions.

Sutton, R. F.↗
NASA NTRS1986

Analysis of separation of a two-stage winged launch vehicle

The feasibility of separating a two-stage winged, parallel configured launch vehicle at Mach 3 has been determined. The two-stage fully-reusable launch vehicle is comprised of an orbiter, which can be used alone as a single-stage-to-orbit vehicle, and a booster which is identical to the baseline orbiter design. A three-degree-of-freedom analysis was conducted on the ascent trajectory of the two-stage concept to determine the optimum staging conditions. The booster provides all propellant for both stages using a crossfeed system until staging occurs. After staging, the empty booster glides back to the launch site for a horizontal landing. To analyze the separation maneuver, the six-degree-of-freedom equations of motion for both stages were numerically integrated using a computer program which incorporates analytically derived interference aerodynamic data. The necessary conditions for a successful separation were determined and required the use of active controls on both the orbiter and booster. Once the booster and orbiter were separated, a three-degree-of-freedom analysis was conducted on the booster's glideback trajectory.

Naftel, J. C.↗
NASA NTRS1997

Effects of Compression, Staging, and Braid Angle on Braided Rope Seal Performance

Future turbine engines and industrial systems will be operating at increased temperatures to achieve more demanding efficiency and performance goals. In the highest temperature sections of the engine new material systems such as ceramics and intermetallics are being considered to withstand the harsh thermal environment. Components constructed of these low expansion-rate materials experience thermal strains and a resulting reduction of life when rigidly attached to high expansion-rate, superalloy support structures. Seals are being designed to both seal and to serve as compliant mounts allowing for relative thermal growths between high temperature but brittle primary structures and the surrounding support structures. Previous seal research yielded several braided rope seal designs which demonstrated the ability to both seal and serve as a compliant mount. The hybrid seal was constructed of an all-ceramic (alumina-silica) core overbraided with a superalloy wire sheath (cobalt based superalloy). The all ceramic seal was constructed of an all-ceramic (alumina-silica) core overbraided with multiple ceramic (alumina-silica) sheath layers. Program goals for braided rope seals are to improve flow resistance and/or seal resilience. To that end, the current report studies the test results of: baseline and modified hybrid seals; two stage hybrid and two stage all-ceramic seal configurations; and single stage hybrid and single stage all-ceramic seal configurations for a range of seal crush conditions. Hybrid seal modifications include increasing the sheath braid angle and core coverage. For the same percent seal cross-sectional crush, results show that increasing the hybrid seal braid angle increased seal stiffness and seal unit load, resulting in flows approximately one third of the baseline hybrid seal flows. For both hybrid and all-ceramic seals, two stage seal configurations significantly outperformed single stage configurations. Two stage seal flows were at least 30% less than the single stage seal flows for the same seal crush. Furthermore, test results of single stage seals indicate that for both all-ceramic and hybrid seals, a specific seal crush condition exists at which minimum flows are achieved (i.e. increasing seal crush beyond a certain point does not result in better flow performance). Flow results are presented for a range of pressures and temperatures from ambient to 1300 F, before and after scrubbing. Compression tests results show that for both all-ceramic and hybrid seals, seal preload and stiffness increase with seal crush, but residual seal interference remains constant.

Steinetz, Bruce M.↗
NASA NTRS2002

The IRIS-GUS Shuttle Borne Upper Stage System

This paper describes the Italian Research Interim Stage - Gyroscopic Upper Stage (IRIS-GUS) upper stage system that will be used to launch NASA's Triana Observatory from the Space Shuttle. Triana is a pathfinder earth science mission being executed on rapid schedule and small budget, therefore the mission's upper stage solution had to be a system that could be fielded quickly at relatively low cost and risk. The building of the IRIS-GUS system wa necessary because NASA lost the capability to launch moderately sized upper stage missions fro the Space Shuttle when the PAM-D system was retired. The IRIS-GUS system restores this capability. The resulting system is a hybrid which mates the existing, flight proven IRIS (Italian Research Interim Stage) airborne support equipment to a new upper stage, the Gyroscopic Upper Stage (GUS) built by the GSFC for Triana. Although a new system, the GUS exploits flight proven hardware and design approaches in most subsystems, in some cases implementing proven design approaches with state-of-the-art electronics. This paper describes the IRIS-GUS upper stage system elements, performance capabilities, and payload interfaces.

Tooley, Craig↗
NASA NTRS2004

Experimental Investigation of a High Pressure Ratio Aspirated Fan Stage

The experimental investigation of an aspirated fan stage designed to achieve a pressure ratio of 3.4:1 at 1500 ft/sec is presented in this paper. The low-energy viscous flow is aspirated from diffusion-limiting locations on the blades and flowpath surfaces of the stage, enabling a very high pressure ratio to be achieved in a single stage. The fan stage performance was mapped at various operating speeds from choke to stall in a compressor facility at fully simulated engine conditions. The experimentally determined stage performance, in terms of pressure ratio and corresponding inlet mass flow rate, was found to be in good agreement with the three-dimensional viscous computational prediction, and in turn close to the design intent. Stage pressure ratios exceeding 3:1 were achieved at design speed, with an aspiration flow fraction of 3.5 percent of the stage inlet mass flow. The experimental performance of the stage at various operating conditions, including detailed flowfield measurements, are presented and discussed in the context of the computational analyses. The sensitivity of the stage performance and operability to reduced aspiration flow rates at design and off design conditions are also discussed.

Merchant, Ali↗
NASA NTRS1957

Performance Evaluation of Reduced-Chord Rotor Blading as Applied to J73 Two-Stage Turbine

The multistage turbine from the J73 turbojet engine has previously been investigated with standard and with reduced-chord rotor blading in order to determine the individual performance characteristics of each configuration over a range of over-all pressure ratio and speed. Because both turbine configurations exhibited peak efficiencies of over 90 percent, and because both units had relatively wide efficient operating ranges, it was considered of interest to determine the performance of the first stage of the turbine as a separate component. Accordingly, the standard-bladed multistage turbine was modified by removing the second-stage rotor disk and stator and altering the flow passage so that the first stage of the unit could be operated independently. The modified single-stage turbine was then operated over a range of stage pressure ratio and speed. The single-stage turbine operated at a peak brake internal efficiency of over 90 percent at an over-all stage pressure ratio of 1.4 and at 90 percent of design equivalent speed. Furthermore, the unit operated at high efficiencies over a relatively wide operating range. When the single-stage results were compared with the multistage results at the design operating point, it was found that the first stage produced approximately half the total multistage-turbine work output.

Schurn, Harold J.↗
NASA NTRS2007

TBCC Fan Stage Operability and Performance

NASA s Fundamental Aeronautics Program is investigating turbine-based propulsion systems for access to space because it provides the potential for aircraft-like, space-launch operations that may significantly reduce launch costs and improve safety. Studies performed under NASA s NGLT and the NASP High Speed Propulsion Assessment (HiSPA) program indicated a variable cycle turbofan/ramjet was the best configuration to satisfy access-to-space mission requirements because this configuration maximizes the engine thrust-to-weight ratio while minimizing frontal area. To this end, NASA and GE teamed to design a Mach 4 variable cycle turbofan/ramjet engine for access to space. To enable the wide operating range of a Mach 4+ variable cycle turbofan ramjet required the development of a unique fan stage design capable of multi-point operation to accommodate variations in bypass ratio (10X), fan speed (7X), inlet mass flow (3.5X), inlet pressure (8X), and inlet temperature (3X). The primary goal of the fan stage was to provide a high pressure ratio level with good efficiency at takeoff through the mid range of engine operation, while avoiding stall and losses at the higher flight Mach numbers, without the use of variable inlet guide vanes. Overall fan performance and operability therefore requires major consideration, as competing goals at different operating points and aeromechanical issues become major drivers in the design. To mitigate risk of meeting the unique design requirements for the fan stage, NASA and GE teamed to design and build a 57% engine scaled fan stage to be tested in NASA s transonic compressor facility. The objectives of this test are to assess the aerodynamic and aero mechanic performance and operability characteristics of the fan stage over the entire range of engine operation including: 1) sea level static take-off, 2) transition over large swings in fan bypass ratio, 3) transition from turbofan to ramjet, and 4) fan windmilling operation at high Mach flight conditions. In addition, the fan stage design was validated by performing pre-test CFD analysis using both GE proprietary and NASA s APNASA codes. Herein we will discuss 1) the fan stage design, 2) the experiment including the unique facility and instrumentation, and 3) the comparison of pre-test CFD analysis to initial aerodynamic test results for the baseline fan stage configuration. Measurements and pre-test analysis will be compared at 37%, 50%, 80%, 90%, and 100% of design speed to assess the ability of state-of-the-art design and analysis tools to meet the fan stage performance and operability requirements for turbine based propulsion for access to space.

Suder, Kenneth L.↗