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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.

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At least 325 records · Page 18
All sourcesNASA NTRSDOE OSTI
NASA NTRS2014

Aerodynamic Effects of High Turbulence Intensity on a Variable-Speed Power-Turbine Blade With Large Incidence and Reynolds Number Variations

The effects of high inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. These results are compared to previous measurements made in a low turbulence environment. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The current study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Assessing the effects of turbulence at these large incidence and Reynolds number variations complements the existing database. Downstream total pressure and exit angle data were acquired for 10 incidence angles ranging from +15.8deg to −51.0deg. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.12×10(exp 5) to 2.12×10(exp 6) and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 8 to 15 percent for the current study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitch/yaw probe located in a survey plane 7 percent axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the turbulence grid imposes.

Flegel, Ashlie B.↗
NASA NTRS2014

Aerodynamic Effects of Turbulence Intensity on a Variable-Speed Power-Turbine Blade with Large Incidence and Reynolds Number Variations

The effects of inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The high turbulence study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Downstream total pressure and exit angle data were acquired for ten incidence angles ranging from +15.8 to 51.0. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.12105 to 2.12106 and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 0.25 - 0.4 for the low Tu tests and 8- 15 for the high Tu study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitchyaw probe located in a survey plane 7 axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the turbulence grid imposes.

Turbines↗
NASA NTRS2014

Aerodynamic Effects of High Turbulence Intensity on a Variable-Speed Power-Turbine Blade with Large Incidence and Reynolds Number Variations

The effects of high inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. These results are compared to previous measurements made in a low turbulence environment. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The current study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Assessing the effects of turbulence at these large incidence and Reynolds number variations complements the existing database. Downstream total pressure and exit angle data were acquired for 10 incidence angles ranging from +15.8deg to −51.0deg. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.12×10(exp 5) to 2.12×10(exp 6) and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 8 to 15 percent for the current study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitch/yaw probe located in a survey plane 7 percent axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the turbulence grid imposes.

Flegel, Ashlie B.↗
NASA NTRS2015

Method for Making Measurements of the Post-Combustion Residence Time in a Gas Turbine Engine

A system and method of measuring a residence time in a gas-turbine engine is provided, whereby the method includes placing pressure sensors at a combustor entrance and at a turbine exit of the gas-turbine engine and measuring a combustor pressure at the combustor entrance and a turbine exit pressure at the turbine exit. The method further includes computing cross-spectrum functions between a combustor pressure sensor signal from the measured combustor pressure and a turbine exit pressure sensor signal from the measured turbine exit pressure, applying a linear curve fit to the cross-spectrum functions, and computing a post-combustion residence time from the linear curve fit.

Miles, Jeffrey H↗
NASA NTRS2017

Method for Making Measurements of the Post-Combustion Residence Time in a Gas Turbine Engine

A method of measuring a residence time in a gas-turbine engine is disclosed that includes measuring a combustor pressure signal at a combustor entrance and a turbine exit pressure signal at a turbine exit. The method further includes computing a cross-spectrum function between the combustor pressure signal and the turbine exit pressure signal, calculating a slope of the cross-spectrum function, shifting the turbine exit pressure signal an amount corresponding to a time delay between the measurement of the combustor pressure signal and the turbine exit pressure signal, and recalculating the slope of the cross-spectrum function until the slope reaches zero.

Miles, Jeffrey H.↗
NASA NTRS2019

Geared Turbofan Engine Having a Reduced Number of Fan Blades and Improved Acoustics

A rotor blade comprises an airfoil extending radially from a root section to a tip section and axially from a leading edge to a trailing edge, the leading and trailing edges defining a curvature therebetween. The curvature determines a relative exit angle at a relative span height between the root section and the tip section, based on an incident flow velocity at the leading edge of the airfoil and a rotational velocity at the relative span height. In operation of the rotor blade, the relative exit angle determines a substantially flat exit pressure ratio profile for relative span heights from 75% to 95%, wherein the exit pressure ratio profile is constant within a tolerance of 10% of a maximum value of the exit pressure ratio profile.

Baltas, Constantine↗
NASA NTRS2019

Design of a Nozzle for the Spyder 2nd Stage Solid Rocket Motor

During the 2019 summer term, the author worked with a group of four interns to complete the preliminary design of a 2nd stage solid rocket motor for Up Aerospace’s Spyder Launch Vehicle. The Spyder vehicle is a four stage, solid fuel rocket designed as part of collaboration between NASA and Up Aerospace to develop a vehicle capable of delivering a 10 kg, 6U CubeSat into a 350 km, circular, low Earth orbit. As part of the agreement, NASA is tasked with designing high performance 2nd, 3rd, and 4th stages while Up Aerospace will provide the 1st stage, derived from the first stage of the company’s sub-orbital Spaceloft XL vehicle. Previous intern teams have designed the 3rd and 4th stages, which left the preliminary design of the 2nd stage motor to be completed this summer. The purpose of this report is to highlight a trade study which the author conducted to determine the nozzle geometry which would most benefit the performance of the 2nd stage motor. In this study, various nozzle parameters such as throat radius (RSI), expansion ratio, mass and their effects on the Isp and Delta V of the 2nd Stage were investigated. From this study, a nozzle geometry providing the necessary performance was chosen and implemented as part of the preliminary design of the 2nd stage motor. To mature the 2nd stage motor design, a trade space was needed to determine the nozzle configuration which would most benefit the performance of the 2nd stage. The trade space established did not only evaluate different expansion ratios for the same throat radius, but also investigated the possible performance gained from decreasing the throat radius to increase the expansion ratio and Isp capable of being delivered by the nozzle. Decreasing the throat radius would cause the chamber pressure to increase, consequently increasing the case and insulation mass required to safely operate a motor at higher pressures. To account for this factor, accurate estimates of inert mass first needed to be established. After doing so, the effects of varying nozzle expansion ratios, exit half angles, and subsequently length and mass were evaluated against motor and nozzle performance factors such as delta V and Isp. For this study, four throat radii ranging from 1.75” to 2.375” and consequently four different chamber pressures ranging 550 psia to 1200 psia were investigated. BACKGROUND To launch into Low Earth Orbit, a payload needs to be accelerated to the orbital velocity necessary to keep it from falling back to Earth. The change in velocity required between launch and orbital insertion is known as Delta V. The Delta V which a rocket or stage can deliver can be calculated using the Ideal Rocket equation, 𝛥𝑉=−𝑔0∗𝐼𝑠𝑝∗ln(𝑀𝑓𝑀𝑖) (3) Where 𝑔0 is the acceleration due to gravity at the earth’s surface, 𝐼𝑠𝑝 is the specific impulse of the rocket, 𝑀𝑖 is the initial mass of the rocket, and 𝑀𝑓 is the final mass of the rocket after burnout. From preliminary calculations beyond the scope of this paper, it was determined that 30500 ft/s of delta V would be required for a payload to be inserted into a 350 km circular orbit around the Earth. Using the known masses and Isp values of the 1st, 3rd, and 4th stages and equation 3, the delta V of each stage was calculated. The delta V required by the 2nd stage could then be found by taking the difference between the total delta V required and the delta V of the 1st, 3rd, and 4th stages. From this, the required delta V of the 2nd Stage was calculated to be 7340 ft/s. Specific impulse is an efficiency factor of the nozzle which defines the impulse delivered by the motor per unit of propellant weight. The main variables of a nozzle’s specific impulse investigated in this trade were exit cone half angle, throat radius, and expansion ratio which is affected by the throat radius. The expansion ratio, ε, of a nozzle is defined as the ratio between the nozzle exit area and throat area, and can be calculated using the equation, ε=𝑅𝑒𝑥𝑖𝑡2𝑅𝑠𝑖2 (2) Where 𝑅𝑒𝑥𝑖𝑡 is the radius of the nozzle’s exit and 𝑅𝑠𝑖 is the radius of the nozzle’s throat. A larger expansion ratio and smaller exit half angle will increase the Isp of a nozzle by allowing the gas to expand more and by allowing more of the exhaust gas to produce thrust in the direction of the motor’s central axis. A cross section view of the 2nd Stage motor with the major components annotated is provided in figure 1.

Bennett, Daniel↗
NASA NTRSDate not supplied

Numerical Study of Combustor-Turbine Interactions Considering a Two-Stage High-Pressure Turbine

In this study, preliminary results are presented from the Open National Combustion Code (OpenNCC) applied to the two-stage high-pressure turbine (HTP) from the energy efficient engine (E3) program. Although the combustor is designed for efficient fuel–air mixing and a relatively uniform exit temperature profile, temporal and spatial variation in combustor exit conditions inevitably lead to hot-streaks in the HPT. These hot streaks can cause local hot spots on the HPT blade surfaces that reduce blade life. To improve understanding of these effects and the overall design of combustor and HPT, we are developing a fully-coupled combustor-turbine computational fluid dynamics (CFD) capability. Previous work has simulated the E^3 combustor at the simulated Sea-Level Take-off (SLTO) condition (P(41) = 2.4 [atm]) with and without the first-stage HPT stator included. This work addresses HPT simulations using inflow conditions based on previous simulation results of the E^3 combustor, as well as investigating the effects of spatial and temporal temperature non-uniformity of the combustor-exit flow on the HPT performance for a more realistic SLTO (P(41) = 27.4 [atm])). First, we analyze existing results from the combustor and the HPT 1st-stage stator simulation, where the unsteady hot gas coming out of the combustor exit travels between the stationary stators and then directly exits the computational domain (i.e., no effect from the rotors). Second, we simulate the HPT including the 1st and 2nd stage stators and rotors with the time-averaged, spatially-nonuniform inflow to the HPT taken from the combustor (with HPT 1st-stage stators) solution. Differences in simulations results for the flow field in the 1st-stage stator region are presented. The results of these two simulations (combustor and HPT simulated in a sequential manner) will be compared to a fully-coupled EEE combustor/HPT simulation to understand the impacts of fully-coupled simulations on combustor and HPT performance for two different operating conditions.

LES↗
NASA NTRSDate not supplied

Numerical Study of Combustor-Turbine Interactions Considering a Two-Stage High-Pressure Turbine

In this study, preliminary results are presented from the Open National Combustion Code (OpenNCC) applied to the two-stage high-pressure turbine (HTP) from the energy efficient engine (E3) program. Although the combustor is designed for efficient fuel–air mixing and a relatively uniform exit temperature profile, temporal and spatial variation in combustor exit conditions inevitably lead to hot-streaks in the HPT. These hot streaks can cause local hot spots on the HPT blade surfaces that reduce blade life. To improve understanding of these effects and the overall design of combustor and HPT, we are developing a fully-coupled combustor-turbine computational fluid dynamics (CFD) capability. Previous work has simulated the E3 combustor at the simulated Sea-Level Take-off (SLTO) condition (P41 = 2.4 [atm]) with and without the first-stage HPT stator included. This work addresses HPT simulations using inflow conditions based on previous simulation results of the E3 combustor, as well as investigating the effects of spatial and temporal temperature non-uniformity of the combustor-exit flow on the HPT performance for a more realistic SLTO (P41 = 27.4 [atm])). First, we analyze existing results from the combustor and the HPT 1st-stage stator simulation, where the unsteady hot gas coming out of the combustor exit travels between the stationary stators and then directly exits the computational domain (i.e., no effect from the rotors). Second, we simulate the HPT including the 1st and 2nd stage stators and rotors with the time-averaged, spatially-nonuniform inflow to the HPT taken from the combustor (with HPT 1st-stage stators) solution. Differences in simulations results for the flow field in the 1st-stage stator region are presented. The results of these two simulations (combustor and HPT simulated in a sequential manner) will be compared to a fully-coupled EEE combustor/HPT simulation to understand the impacts of fully-coupled simulations on combustor and HPT performance for two different operating conditions.

LES↗
DOE OSTI2026

Investigation of the 244 Pu ⁢( 48 Ca,𝑥⁢𝑛) 292−𝑥 Fl reaction with the LBNL SHREC detector: Investigation of decay chains of isotopes of flerovium (𝑍=114)

The 244 Pu ⁢( 48 Ca,𝑥⁢𝑛)⁢ 292−𝑥 Fl reaction was investigated at Lawrence Berkeley National Laboratory’s 88 Inch Cyclotron using the Berkeley Gas-filled Separator (BGS), the newly installed Superheavy Recoil detector, along with an upgraded digital electronics and data acquisition system. Seven decay chains were observed starting with an evaporation residue, followed by a single 𝛼 decay and a spontaneous fission. The decay characteristics of these seven decay chains led to an assignment to 288 Fl , the product of the 4⁢𝑛 reaction channel. Two additional chains were (tentatively) assigned to the decay of the 3⁢𝑛 exit channel, 289 Fl . Cross sections for the 4⁢𝑛 and 3⁢𝑛 exit channels were 𝜎 prod =6.7⁢($^{36}_{25}$) pb and 𝜎 prod =1.6⁢($^{22}_{11}$) pb, respectively. Another decay chain, tentatively assigned to the 5⁢𝑛 exit channel through the 48 Ca + 244 Pu reaction or the 3⁢𝑛 exit channel of the 48 Ca + 242 Pu reaction, was also detected. Detailed information regarding the observed decay chains and their nuclear structure aspects is discussed, along with the performance of the BGS and the new detection system.

73 NUCLEAR PHYSICS AND RADIATION PHYSICS↗
NASA NTRS1972

Computer program for preliminary design analysis of axial-flow turbines

The program method is based on a mean-diameter flow analysis. Input design requirements include power or pressure ratio, flow, temperature, pressure, and speed. Turbine designs are generated for any specified number of stages and for any of three types of velocity diagrams (symmetrical, zero exit swirl, or impulse). Exit turning vanes can be included in the design. Program output includes inlet and exit annulus dimensions, exit temperature and pressure, total and static efficiencies, blading angles, and last-stage critical velocity ratios. The report presents the analysis method, a description of input and output with sample cases, and the program listing.

Glassman, A. J.↗
NASA NTRS1972

Normal impingement loads due to small air jets issuing from a base plate and reflecting off a platform for various jet Mach numbers, separation distances, and ambient pressures

An investigation was conducted in a 12.5-meter-diameter vacuum sphere to determine the impingement loads due to air jets issuing from and perpendicular to a circular base and reflecting off a square platform, that is, a simulation of rendezvous maneuvering, docking, launch, impact dampers etc. The nozzles had exit Mach numbers of 1, 3, 5, and 7. The ambient pressures were 0.0006, 5, 225, and 760 torr. Under near-field separation distances and at 0.0006 torr, reflections were significant; and ratios of the impingement force to thrust on both plates in the biplane arrangement varied from about 750 for exit Mach number 1 to 120 for exit Mach number 7. The far-field force ratios were near unity for the platform and zero for the base and indicated few, if any, reflections. Some reversals and rapid changes in loads were obtained at transition distances between the near and far fields. In general, increasing the exit Mach number or ambient pressure reduced the impingement loads.

Hoffman, S.↗
NASA NTRS1974

Boundary-layer analysis of subsonic inlet diffuser geometries for engines nacelles

Theoretical Mach number distributions and boundary-layer parameters are presented for subsonic nacelle inlet diffuser geometries with length to exit diameter ratios ranging from 0.4 to 1.6 and diffuser exit area to throat area ratios ranging from 1.1 to 2.0. The major portion of the study was done with a cubic diffuser contour with the inflection point at the midpoint of the diffuser, a diffuser throat Mach number of 0.6, and a free-stream Mach number of 0.12. Calculations were performed at both model (diffuser exit diameter, 30.5 cm) and full-scale (diffuser exit diameter, 183 cm) sizes. Separation limits were defined by establishing a separation boundary on plots of diffuser area ratio as a function of diffuser length to diameter ratio. The effects of diffuser contour, inlet lip geometry, and throat Mach number on the boundary-layer characteristics are illustrated. The major results of the study indicate that the separation boundary is shifted to greater area ratios by (1) increasing the diffuser length, (2) increasing the scale of the diffuser and, (3) moving the inflection point of the diffuser contour to or ahead of the midpoint of the diffuser.

Albers, J. A.↗
NASA NTRS1974

Research on free and impinging jets for the development of STOL aircraft

The effect of the exit plane conditions on the initial region of an axisymmetric jet was systematically investigated. An essentially top-hat mean velocity profile and a homogeneous turbulence structure were maintained at the exit plane for eight distinct scale and intensity conditions. Mass and momentum flux values are independent of the exit turbulence structure for the range investigated; however, a significant (approximately 25%) increase in the latter implies a pronounced static pressure decrement inside the jet. Details of the velocity profile and turbulence structure are influenced by the exit plane conditions. The three radial-axial components of the Reynolds stress tensor have been conditionally sampled and are analyzed to show the initial condition effects.

Foss, J. F.↗
NASA NTRS1972

Jet engine exhaust noise due to rough combustion and nonsteady aerodynamic sources

Internal sources are accounted for in terms of fluctuations of mass and momentum at the nozzle exit plane. At low Mach numbers, mass flow fluctuations generated at the exit plane by acoustic resonant type fluctuations inside the engine are found to be dominant. In the subsonic Mach number range between 0.3 and 0.5, exit plane mass flow fluctuations at frequencies characteristic of turbulence become most dominant. Above Mach 0.5, the turbulent momentum fluctuations at the exit plane become dominant, and the jet contribution is not found significant at subsonic speeds.

Plett, E. G.↗
NASA NTRS1975

Aerodynamic performance of a fully film cooled core turbine vane tested with cold air in a two-dimensional cascade

The aerodynamic performance of a fully film cooled core turbine vane was investigated experimentally in a two-dimensional cascade of 10 vanes. Three of the 10 vanes were cooled; the others were solid (uncooled) vanes. Cold air was used for both the primary and coolant flows. The cascade test covered a range of pressure ratios corresponding to ideal exit critical velocity ratios of 0.6 to 0.95 and a range of coolant flow rates to 7.5 percent of the primary flow. The coolant flow was varied by changing the coolant supply pressure. The principal measurements were cross-channel surveys of exit total pressure, static pressure, and flow angle. The results presented include exit survey data and overall performance in terms of loss, flow angle, and weight flow for the range of exit velocity ratios and coolant flows investigated. The performance of the cooled vane is compared with the performance of an uncooled vane of the same profile and also with the performance obtained with a single cooled vane in the 10-vane cascade.

Stabe, R. G.↗
NASA NTRS1977

Flow characteristics in an expansion tunnel as inferred from velocity measurements

A photoionization technique was used to study flow characteristics in an expansion tunnel. Vertical surveys of the axial component of flow velocity just downstream from the nozzle exit were obtained, and estimates of freestream density were inferred from the velocity measurement technique. The pitot pressure was measured and compared to the average axial component of velocity as a function of time for the two cases when air and CO2 were used as test gases. Vertical velocity and static density profiles at the nozzle exit are presented for the case when CO2 was used as test gas. Experimental results were used to determine the diameter and uniformity of the test core at the nozzle exit and the duration of the quasi-steady flow period. These data are relevant to evaluation of the suitability of operating an expansion tube as an expansion tunnel. The expansion tunnel is an expansion tube with a conical nozzle positioned at the exit of the acceleration section, so that nozzle entrance flow conditions are hypersonic and characterized by hypervelocity.

Miller, C. G.↗
NASA NTRS1977

Experimental results of large-scale structures in jet flows and their relation to jet noise production

Experiments have been performed to determine the role of large-scale turbulent structures in the production of jet noise. Axisymmetric turbulent jet flows at ambient stagnation temperature have been observed with the aid of flow visualization techniques. Jet Mach numbers at the nozzle exit ranged between 0.1 and 0.9, and the Reynolds number, based on nozzle exit diameter, was approximately 10 to the 6th. Large organized turbulent structures existed as far downstream of the nozzle exit as 7 diameters. High-speed schlieren motion pictures synchronized with near-field pressure measurements of an excited jet indicated that strong instantaneous peaks in the pressure signal occurred whenever a merging process between two large-scale organized structures occurred. This pressure pulse propagated at a speed which was somewhat larger than the velocity of the jet at the nozzle exit.

Sarohia, V.↗