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DOE OSTI2023

Effects of additively manufactured surface roughness on small diameter plenums with multiple side discharges for high temperature gas turbine blade cooling applications

Combined heat and power (CHP) applications have significant environmental and economic benefits that are consistent with the goals of the U.S. Department of Energy (DOE). One area that is currently being studied includes the potential benefits of CHP turbine operation at higher turbine inlet temperatures. Internal cooling concepts enabled by additive manufacturing (AM) are of primary interest. Here, the effectiveness of internal cooling is hindered by many factors such as velocity distribution of the cooling air to the hot surface considering impingement cooling. To simulate cooling air exiting from a series of orifices for internal cooling in an airfoil, a straight smooth wall tubing with multiple side discharging orifices is used and compared to additively manufactured tubing (Ti6Al4V Grade 23) with orifice size and spacing as well as inner and outer diameters identical to the smooth wall tubing. Similar to flow discharging form perforated pipes, the flow discharged from individual orifices along the tubes in this study is found to be nonuniformly distributed, and the horizontal (axial direction) momentum can be observed from the experimental data. Discharge velocities have been measured with two-dimensional (2D) particle imaging velocimetry (PIV) and single element hot wire anemometry. Numerical analysis has also been conducted to predict the velocity distributions along the orifices in the smooth wall and additive manufacturing (AM) tubing, which are inherited with surface roughness. Numerical and measured results in this study are compared, presented, and discussed.

32 ENERGY CONSERVATION, CONSUMPTION, AND UTILIZATI↗

NASA NTRS2006

An Advanced Helium Buffer Seal for the SSME, ATD Oxygen Pump

The present configuration of Helium Buffer Seal on the ATD oxygen pump consists of a pair of opposed carbon rings are forced axially against their containment housings. Leakage occurs through the clearance between the rings and the shaft. The total helium leakage through both sides is approximately 239 SCFM. A reduction in leakage to 50 SCFM will result in less helium storage and consequently permit a substantial increase in payload. Under Phase 1 NASA SBIR, a solid T-Ring seal was analyzed and designed that could satisfy the criteria of reducing leakage to 50 SCFM or less. The design makes maximum use of available length and employs a mid length row of hydrostatic orifaces that feed buffer helium directly into a 2 to 3 mil clearance region. The flow splits into opposite paths to buffer oxygen gas on one side and hydrogen gas on the turbine side. The seal employs opposed hydrostatic tapered land secondary seals that provide friction free support of the primary seal and allows the primary seal to follow rotor excursion and maintain concentric operating clearance . The predicted performance of the T-seal is excellent with operation at a safe film thickness of 2 to 2.5 mils and leakage less than 50 SCFM.

Shapiro, Wilbur↗

NASA NTRS1999

Simulations of the Unsteady Flow through the Fastrac Supersonic Turbine

Analysis of the unsteady aerodynamic environment in the Fastrac supersonic turbine is presented. Modal analysis of the turbine blades indicated possible resonance in crucial operating ranges of the turbopump. Unsteady computational fluid dynamics (CFD) analysis was conducted to support the aerodynamic and structural dynamic assessments of the turbine. Before beginning the analysis, two major problems with current unsteady analytical capabilities had to be addressed: modeling a straight centerline nozzle with the turbine blades and exit guide vanes (EGVs), and reducing run times significantly while maintaining physical accuracy. Modifications were made to the CFD code used in this study to allow the coupled nozzle/blade/EGV analysis and to incorporate Message Passing Interface (MPI) software. Because unsteadiness is a key issue for the Fastrac turbine [and future rocket engine turbines such as for the Reusable Launch Vehicle (RI.V)], calculations were performed for two nozzle-to-blade axial gaps. Calculations were also performed for the nozzle alone, and the results were imposed as an inlet boundary condition for a blade/EGV calculation for the large gap case. These results are compared to the nozzle/blade/EGV results.

Griffin, Lisa W.↗

NASA NTRS1999

Water Flow Performance of a Superscale Model of the Fastrac Liquid Oxygen Pump

As part of the National Aeronautics and Space Administration's ongoing effort to lower the cost of access to space, the Marshall Space Flight Center has developed a rocket engine with 60,000 pounds of thrust for use on the Reusable Launch Vehicle technology demonstrator slated for launch in 2000. This gas generator cycle engine, known as the Fastrac engine, uses liquid oxygen and RP-1 for propellants and includes single stage liquid oxygen and RP-1 pumps and a single stage supersonic turbine on a common shaft. The turbopump design effort included the first use and application of new suction capability prediction codes and three-dimensional blade generation codes in an attempt to reduce the turbomachinery design and certification costs typically associated with rocket engine development. To verify the pump's predicted cavitation performance, a water flow test of a superscale model of the Fastrac liquid oxygen pump was conducted to experimentally evaluate the liquid oxygen pump's performance at and around the design point. The water flow test article replicated the flow path of the Fastrac liquid oxygen pump in a 1.582x scale model, including scaled seal clearances for correct leakage flow at a model operating speed of 5000 revolutions per minute. Flow entered the 3-blade axial-flow inducer, transitioned to a shrouded, 6-blade radial impeller, and discharged into a vaneless radial diffuser and collection volute. The test article included approximately 50 total and static pressure measurement locations as well as flush-mounted, high frequency pressure transducers for complete mapping of the pressure environment. The primary objectives of the water flow test were to measure the steady-state and dynamic pressure environment of the liquid oxygen pump versus flow coefficient, suction specific speed, and back face leakage flow rate. Results showed excellent correlation between the predicted and experimentally measured pump head rise at low suction specific speeds. Likewise, only small circumferential variations in steady-state impeller exit and radial diffuser pressure distributions were observed from 80% to 120% of the design flow coefficient, matching the computational predictions and confirming that the integrated design approach has minimized any exit volute-induced distortions. The test article exhibited suction performance trends typically observed in inducer designs with virtually constant head rise with decreasing inlet pressure until complete pump head breakdown. Unfortunately, the net positive suction head at 3% head fall-off occurred far below that predicted at all tested flow coefficients, resulting in a negative net positive suction head margin at the design point in water. Additional testing to map the unsteady pressure environment was conducted and interesting rotating phenomena at the inducer inlet were observed. These rotating phenomena's cell numbers, direction, and speed were correlated with pump operating parameters. The impact of the unsteady phenomena and their corresponding energy losses on the unexpectedly poor pump performance is also discussed.

Skelley, Stephen↗

NASA NTRS2001

Water Flow Performance of a Superscale Model of the Fastrac Liquid Oxygen Pump

As part of the National Aeronautics and Space Administration's ongoing effort to lower the cost of access to space, the Marshall Space Flight Center has developed a rocket engine with 60,000 pounds of thrust for use on the Reusable Launch Vehicle technology demonstrator slated for launch in 2000. This gas generator cycle engine, known as the Fastrac engine, uses liquid oxygen and RP-1 for propellants and includes single stage liquid oxygen and RP-1 pumps and a single stage supersonic turbine on a common shaft. The turbopump design effort included the first use and application of new suction capability prediction codes and three-dimensional blade generation codes in an attempt to reduce the turbomachinery design and certification costs typically associated with rocket engine development. To verify the pump's predicted cavitation performance, a water flow test of a superscale model of the Fastrac liquid oxygen pump was conducted to experimentally evaluate the liquid oxygen pump's performance at and around the design point. The water flow test article replicated the flow path of the Fastrac liquid oxygen pump in a 1.582x scale model, including scaled seal clearances for correct leakage flow at a model operating speed of 5000 revolutions per minute. Flow entered the 3-blade axial-flow inducer, transitioned to a shrouded, 6- blade radial impeller, and discharged into a vaneless radial diffuser and collection volute. The test article included approximately 50 total and static pressure measurement locations as well as flush-mounted, high frequency pressure transducers for complete mapping of the pressure environment. The primary objectives of the water flow test were to measure the steady-state and dynamic pressure environment of the liquid oxygen pump versus flow coefficient, suction specific speed, and back face leakage flow rate. Initial results showed acceptable correlation between the predicted and experimentally measured pump head rise at low suction specific speeds. Likewise, only small circumferential variations in steady-state were observed from 80% to 120% of the design flow coefficient, matching the computational predictions and confirming that the integrated design approach has minimized any exit volute-induced distortions. The test article exhibited suction performance trends typically observed in inducer designs with virtually constant head rise with decreasing inlet pressure until complete pump head breakdown. Unfortunately, the net positive suction head at 3% head fall-off occurred far below that predicted at all tested flow coefficients, resulting in a negative net positive suction head margin at the design point in water. Additional testing to map the unsteady pressure environment was conducted and cavitation-induced flow disturbances at the inducer inlet were observed. Two distinct disturbances were identified, one rotating and one stationary relative to the fixed frame of reference, while the transition from one regime to the next produced significant effects on the steady state pump performance. The impact of the unsteady phenomena and the corresponding energy losses on the unexpectedly poor pump performance is also discussed.

Skelley, Stephen↗

NASA NTRS1989

An interactive grid generation procedure for axial and radial flow turbomachinery

A combination algebraic/elliptic technique is presented for the generation of three dimensional grids about turbo-machinery blade rows for both axial and radial flow machinery. The technique is built around use of an advanced engineering workstation to construct several two dimensional grids interactively on predetermined blade-to-blade surfaces. A three dimensional grid is generated by interpolating these surface grids onto an axisymmetric grid. On each blade-to-blade surface, a grid is created using algebraic techniques near the blade to control orthogonality within the boundary layer region and elliptic techniques in the mid-passage to achieve smoothness. The interactive definition of bezier curves as internal boundaries is the key to simple construction. This procedure lends itself well to zonal grid construction, an important example being the tip clearance region. Calculations done to date include a space shuttle main engine turbopump blade, a radial inflow turbine blade, and the first stator of the United Technologies Research Center large scale rotating rig. A finite Navier-Stokes solver was used in each case.

Beach, Timothy A.↗

NASA NTRS1990

An interactive grid generation procedure for axial and radial flow turbomachinery

Beach, T. A.↗

NASA NTRS2010

Axial flow positive displacement worm compressor

An axial flow positive displacement compressor has an inlet axially spaced apart and upstream from an outlet. Inner and outer bodies have offset inner and outer axes extend from the inlet to the outlet through first and second sections of a compressor assembly in serial downstream flow relationship. At least one of the bodies is rotatable about its axis. The inner and outer bodies have intermeshed inner and outer helical blades wound about the inner and outer axes respectively. The inner and outer helical blades extend radially outwardly and inwardly respectively. The helical blades have first and second twist slopes in the first and second sections respectively. The first twist slopes are less than the second twist slopes. An engine including the compressor has in downstream serial flow relationship from the compressor a combustor and a high pressure turbine drivingly connected to the compressor by a high pressure shaft.

Murrow, Kurt David↗

NASA NTRS1995

Water driven turbine/brush pipe cleaner

Assemblies are disclosed for cleaning the inside walls of pipes and tubes. A first embodiment includes a small turbine with angled blades axially mounted on one end of a standoff support. An O-ring for stabilizing the assembly within the pipe is mounted in a groove within the outer ring. A replaceable circular brush is fixedly mounted on the opposite end of the standoff support and can be used for cleaning tubes and pipes of various diameters, lengths and configurations. The turbine, standoff support, and brush spin in unison relative to a hub bearing that is fixedly attached to a wire upstream of the assembly. The nonrotating wire is for retaining the assembly in tension and enabling return of the assembly to the pipe entrance. The assembly is initially placed in the pipe or tube to be cleaned. A pressurized water or solution source is provided at a required flow-rate to propel the assembly through the pipe or tube. The upstream water pressure propels and spins the turbine, standoff support and brush. The rotating brush combined with the solution cleans the inside of the pipe. The solution flows out of the other end of the pipe with the brush rotation controlled by the flow-rate. A second embodiment is similar to the first embodiment but instead includes a circular shaped brush with ring backing mounted in the groove of the exterior ring of the turbine, and also reduces the size of the standoff support or eliminates the standoff support.

Werlink, Rudy J.↗

NASA NTRS1983

The performance of an annular vane swirler

In connection with the desirability of optimizing the design of a gas turbine combustion chamber, there exists a need for a more complete understanding of the fluid dynamics of the flow in such chambers. In order to satisfy this need, experimental and theoretical research is being conducted with the objective to study two-dimensional axisymmetric geometries under low speed, nonreacting, turbulent, swirling flow conditions. The flow enters the test section and proceeds into a larger chamber. Inlet swirl vanes are adjustable to a variety of vane angles. The present investigation concentrates on the time-mean flow characteristics which are generated by the upstream annular swirler. The investigation makes use of a five-hole pitot probe technique. A theoretical analysis of swirl numbers associated with several idealized exit velocity profiles is included, and values of the ratio of maximum swirl velocity to maximum axial velocity at different swirl strengths are given for each case.

Lilley, D. G.↗

NASA NTRSDate not supplied

Combustor Emissions Modeling of Sustainable Fuels for Commercial Supersonics Technology

An overview is provided of a CFD assessment with Jet-A replacement fuels and advanced fuel blends to assess fuel impacts on NOx emissions and flame structure in an axially-staged combustor for NASA’s Commercial Supersonic Transport (CST) program. The National Combustor Code (OpenNCC) was used to perform two-phase reacting flow computations with two replacement fuels for Jet-A, and various blending ratios of ‘average’ Jet-A (A2) and GEVO-ATJ (C1), for UTRC’s Axially Controlled Stoichiometry (ACS) combustor at CST cruise conditions. The predicted flame structures in the ACS combustor with for the replacement fuels, and for three different blending ratios of A2 and C1 fuel were very similar to each other. The predicted NOx emissions for the two replacement fuels, and three blending ratios of A2 and C1 were within 15% of the experimentally measured NOx emissions for 100% A2 fuel.

Gas Turbine Combustion↗

NASA NTRS1999

Mixing of Multiple Jets with a Confined Subsonic Crossflow: Opposed Rows of Orifices in Rectangular Ducts - Part 2

This paper summarizes experimental and computational results on the mixing of opposed rows of jets with a confined subsonic crossflow in rectangular ducts. The studies from which these results were excerpted investigated flow and geometric variations typical of the complex three-dimensional flowfield in the combustion chambers in gas turbine engines. The principal observation was that the momentum-flux ratio, J, and the orifice spacing, S/H, were the most significant flow and geometric variables. Jet penetration was critical, and penetration decreased as either momentum-flux ratio or orifice spacing decreased. It also appeared that jet penetration remained similar with variations in orifice size, shape, spacing, and momentum-flux ratio when the orifice spacing was inversely proportional to the square-root of the momentum-flux ratio. It was also seen that planar averages must be considered in context with the distributions. Note also that the mass-flow ratios and the orifices investigated were often very large (jet-to-mainstream mass-flow ratio > 1 and the ratio of orifices-area-to-mainstream- cross-sectional-area up to 0.5, respectively), and the axial planes of interest were often just downstream of the orifice trailing edge. Three-dimensional flow was a key part of efficient mixing and was observed for all configurations.

Holdeman, J. D.↗

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 NTRS1997

Mixing of Multiple Jets With a Confined Subsonic Crossflow: Opposed Rows of Orifices in Rectangular Ducts - Part 2

This paper summarizes experimental and computational results on the mixing of opposed rows of jets with a confined subsonic crossflow in rectangular ducts. The studies from which these results were excerpted investigated flow and geometric variations typical of the complex 3-D flowfield in the combustion chambers in gas turbine engines. The principal observation was that the momentum-flux ratio, J, and the orifice spacing, S/H, were the most significant flow and geometric variables. Jet penetration was critical, and penetration decreased as either momentum-flux ratio or orifice spacing decreased. It also appeared that jet penetration remained similar with variations in orifice size, shape, spacing, and momentum-flux ratio when the orifice spacing was inversely proportional to the square-root of the momentum-flux ratio. It was also seen that planar averages must be considered in context with the distributions. Note also that the mass-flow ratios and the offices investigated were often very large (jet-to-mainstream mass-flow ratio greater than 1 and the ratio of orifices-area-to-mainstream-cross-sectional-area up to 0.5 respectively), and the axial planes of interest were often just downstream of the orifice trailing edge. Three-dimensional flow was a key part of efficient mixing and was observed for all configurations.

Holdeman, James D.↗

DOE OSTI2021

Role of Inlet Boundary Conditions on Fuel-Air Mixing at Supercritical Conditions

Advanced engine design and alternative fuels present the possibility of fuel injection at purely supercritical conditions in diesel engines and gas turbines. The complex interactions that govern this phenomenon still need significant research, particularly the boundary conditions for fuel injection are critical for accurate simulation. However, the flow inside the injector itself is often omitted to reduce the computational efforts, and thus, velocity, mass flux, or total pressure is specified at the injector exit (or domain inlet), often with simplified velocity profiles and turbulence levels. This simplified inlet boundary treatment has minimal effects on results for conventional fuel injection conditions, however, the validity of this approach at supercritical conditions has not been assessed. Comprehensive real-gas and binary fluid mixing models have been implemented for computational fluid dynamics (CFD) analysis of fuel-air mixing at supercritical conditions. The model is verified using prior CFD results from the literature. The model is used to investigate the effects of the shape of axial velocity and mass fraction profiles at the inlet boundary with the goal to improve the comparison of predictions to experimental data. Results show that the boundary conditions have a significant effect on the predictions, and none of the cases match precisely with experimental data. Furthermore, the study reveals that the physical location of the inlet boundary might be difficult to infer correctly from the experiments and highlights the need for high-quality, repeatable measurements at supercritical conditions to support the development of relevant high-fidelity models for fuel-air mixing.

02 PETROLEUM↗

NASA NTRS1996

Mixing of Multiple Jets with a Confined Subsonic Crossflow in a Cylindrical Duct

This paper summarizes NASA-supported experimental and computational results on the mixing of a row of jets with a confined subsonic crossflow in a cylindrical duct. The studies from which these results were derived investigated flow and geometric variations typical of the complex 3-D flowfield in the combustion chambers in gas turbine engines. The principal observations were that the momentum-flux ratio and the number of orifices were significant variables. Jet penetration was critical, and jet penetration decreased as either the number of orifices increased or the momentum-flux ratio decreased. It also appeared that jet penetration remained similar with variations in orifice size, shape, spacing, and momentum-flux ratio when the number of orifices was proportional to the square-root of the momentum-flux ratio. In the cylindrical geometry, planar variances are very sensitive to events in the near wall region, so planar averages must be considered in context with the distributions. The mass-flow ratios and orifices investigated were often very large (mass-flow ratio greater than 1 and ratio of orifice area-to-mainstream cross-sectional area up to 0.5), and the axial planes of interest were sometimes near the orifice trailing edge. Three-dimensional flow was a key part of efficient mixing and was observed for all configurations. The results shown also seem to indicate that non-reacting dimensionless scalar profiles can emulate the reacting flow equivalence ratio distribution reasonably well. The results cited suggest that further study may not necessarily lead to a universal 'rule of thumb' for mixer design for lowest emissions, because optimization will likely require an assessment for a specific application.

Holdeman, James D.↗

NASA NTRS1998

Mixing of Multiple Jets With a Confined Subsonic Crossflow

Results from a recently completed enhanced mixing program are summarized in the two technical papers. These studies were parts of a High Speed Research (HSR)-supported joint Government/industry/university program that involved, in addition to the NASA Lewis Research Center, researchers at United Technologies Research Center, Allison Engine Company, CFD Research Corporation, and the University of California, Irvine. The studies investigated the mixing of jets injected normal to a confined subsonic mainsteam in both rectangular and cylindrical ducts. Experimental and computational studies were performed in both nonreacting and reacting flows. The orifice geometries and flow conditions were selected as typical of the complex three-dimensional flows in the combustion chambers in low-emission gas turbine engines. The principal conclusion from both the experiments and modeling was that the momentum-flux ratio J and orifice spacing S/H were the most significant flow and geometry variables, respectively. Conserved scalar distributions were similar-independent of reaction, orifice diameter H/d, and shape-when the orifice spacing and the square root of the momentum-flux ratio were inversely proportional. Jet penetration was critical, and penetration decreased as either momentum-flux ratio or orifice spacing decreased. We found that planar averages must be considered in context with the distributions. The mass-flow ratios and the orifices investigated were often very large. The jet-to-mainstream mass-flow ratio was varied from significantly less than 1 to greater than 1. The orifice-area to mainstream-cross-sectional-area was varied from approx. 0 to 0.5, and the axial planes of interest were often just downstream of the orifice trailing edge. Three-dimensional flow was a key part of efficient mixing and was observed for all configurations. As an example of the results, the accompanying figure shows the effects of different rates of mass addition on the opposite walls of a rectangular duct.

Holdeman, James D.↗

NASA NTRS1993

Compliant seal development

The compliant metallic seal combines the noncontact feature of the labyrinth seal, the low leakage of a mechanical seal, and the compliant nature of the brush seal. It consists of several thin metallic elements or leaves mounted within a ring which is press fit into the housing, and in form, sort of resembles a lip seal sections wiping the shaft. A second set of overlapping cover leaves are placed on top of the shaft riding leaves which reduces leakage and provides stiffness. The leaves can be straight or angle cut. The shaft riding fingers are designed with mismatched curvature to provide lift off similar to the Rayleigh lift pads in mechanical seals with leading edge clearances nearly twice those of the trailing edge as as shown by Fleming to be optimal for gas flows in convergent seal passages. Leading edge clearances range from 300 to 500 microinches. Balance pockets beneath the leaves provide fluid film feed to the 'Rayleigh lift' surface and the proper balance ratio (mechanical seal) when combined with the static pressure and film pressure. The leaves flex in the radial direction and accommodate thermomechanical behavior as well as axial motion and angular misalignment. In the static mode, there is a net closing force on the leaves. The seals were tested to 70 psi at speeds to 16,000 rpm or surface speeds to 330 fps and temperatures from ambient to 440 F. A slow cycle through the rig critical at 10,000 rpm induced a radial vibration response of 0.004 to 0.005 inch were accommodated by the seal. Preliminary performance data are encouraging demonstrating hydrodynamic liftoff and noncontacting operation at pressure and speeds typical of gas turbine engines. The leakage performance data are significantly better than commercial labyrinth and brush seals which should be expected as this design incorporates the features of the low leakage face or mechanical seal along with the flexibility of the brush configuration.

Hendricks, Robert C.↗