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At least 73 records · Page 4

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NASA NTRS2008

Validation of High-Fidelity CFD Simulations for Rocket Injector Design

Computational fluid dynamics (CFD) has the potential to improve the historical rocket injector design process by evaluating the sensitivity of performance and injector-driven thermal environments to the details of the injector geometry and key operational parameters. Methodical verification and validation efforts on a range of coaxial injector elements have shown the current production CFD capability must be improved in order to quantitatively impact the injector design process. This paper documents the status of a focused effort to compare and understand the predictive capabilities and computational requirements of a range of CFD methodologies on a set of single element injector model problems. The steady Reynolds-Average Navier-Stokes (RANS), unsteady Reynolds-Average Navier-Stokes (URANS) and three different approaches using the Large Eddy Simulation (LES) technique were used to simulate the initial model problem, a single element coaxial injector using gaseous oxygen and gaseous hydrogen propellants. While one high-fidelity LES result matches the experimental combustion chamber wall heat flux very well, there is no monotonic convergence to the data with increasing computational tool fidelity. Systematic evaluation of key flow field regions such as the flame zone, the head end recirculation zone and the downstream near wall zone has shed significant, though as of yet incomplete, light on the complex, underlying causes for the performance level of each technique. 1 Aerospace Engineer and Combustion CFD Team Leader, MS ER42, NASA MSFC, AL 35812, Senior Member, AIAA. 2 Professor and Director, Computational Combustion Laboratory, School of Aerospace Engineering, 270 Ferst Dr., Atlanta, GA 30332, Associate Fellow, AIAA. 3 Reilly Professor of Engineering, School of Mechanical Engineering, 585 Purdue Mall, West Lafayette, IN 47907, Fellow, AIAA. 4 Principal Member of Technical Staff, Combustion Research Facility, 7011 East Avenue, MS9051, Livermore, CA 94550, Associate Fellow, AIAA. 5 J. L. and G. H. McCain Endowed Chair, Mechanical Engineering, 104 Research Building East, University Park, PA 16802, Fellow, AIAA. American Institute of Aeronautics and Astronautics 1

Tucker, P. Kevin↗

NASA NTRS2015

Gas Turbine Engine Staged Fuel Injection Using Adjacent Bluff Body and Swirler Fuel Injectors

A fuel injection array for a gas turbine engine includes a plurality of bluff body injectors and a plurality of swirler injectors. A control operates the plurality of bluff body injectors and swirler injectors such that bluff body injectors are utilized without all of the swirler injectors at least at low power operation. The swirler injectors are utilized at higher power operation.

Snyder, Timothy S.↗

NASA NTRS2018

Comparison of Combustion Dynamic Characteristics of Two Advanced Multi-Cup Fuel Injectors

An experimental investigation of the combustion dynamic characteristics of two advanced multi-cup lean direct injectors (LDI) under simulated gas turbine combustor conditions was conducted. The objective was to gain a better understanding of the physical phenomena inside a pressurized flame tube combustion chamber and study the effects of injector flow number on combustion dynamics. The injectors are known as Three-zone Injectors one and two or 3ZI-1 and 3ZI-2, respectively. The injectors were experimentally evaluated at inlet pressures up to 1.724 MPa, non-vitiated air temperatures up to 828K, and adiabatic flame temperatures up to 1975K. Dynamic pressure measurements were taken upstream of the injectors and in the combustion zone. The combustion dynamic behavior of the two injectors was measured over a range of inlet pressures, inlet temperatures, fuel air ratios, and fuel flow splits.

Combustion Dynamics↗

NASA NTRSDate not supplied

Numerical Modeling of No Vent Filling of a Cryogenic Tank with Thermo-dynamic Vent System Assisted Injector

This paper presents a multi-node finite volume model of No Vent Filling (NVF) of a cryogenic tank with Thermo-dynamic Vent System (TVS) assisted injector, using the Generalized Fluid System Simulation Program (GFSSP), a general purpose flow network code. NVF tests were conducted in a CRYOgenic Orbital Testbed (CRYOTE) tank which was filled by liquid nitrogen using a TVS assisted injector with vent valve closed during the entire filling process. In TVS assisted injector, the liquid flow splits into two streams: one stream is routed to a Joule-Thomson (J-T) orifice where the flow immediately flashes from liquid to vapor or two phase mixture because the downstream to J-T leg is maintained at vacuum level.; the other stream is injected into the tank after being cooled by cold vapor of J-T leg in a heat exchanger. The flow through J-T leg is also used to cool the outer metal matrix of the injector which in turn cools the vapor in the tank ullage. The cooling of vapor in the ullage by cold injector surface reduces ullage pressure that allows liquid to enter and fill the tank. An integrated numerical model of the test set up was developed. The model included the dual lines of TVS assisted injector, tank and tank wall. The tank was discretized into multiple fluid nodes and branches to represent the ullage and liquid nitrogen and multiple solid nodes to represent the tank wall and structure. The heat transfer between solid to fluid was calculated from pool boiling correlations which include film, transition, and nucleate boiling, as well as natural convection during pre- and post-boiling. The model also accounts for the condensation of vapor at ullage-injector interface and when it comes in contact with the liquid spray. The predicted pressure, resident mass, wall temperature in the tank were compared with the test data.

Alok Majumdar↗

NASA NTRSDate not supplied

A Review Towards the Design Optimization of High-Performance Additively Manufactured Rotating Detonation Rocket Engine Injectors

Rotating Detonation Rocket Engines (RDRE) have been marketed primarily for their higher specific impulse potential over constant pressure (CP) liquid rocket engines. However, several other performance advantages exist such as heat transfer advantages for gas expander cycle, increased completeness of combustion at low chamber L*, compact engine design, reduced coolant channel pressure drop potential, and improved injector C* performance. NASA has paved the way for liquid engine system performance enhancement since the Apollo program and continues to do so with metal additive manufacturing (AM), new advanced materials, and advanced propulsion concepts. A team of propulsion development engineers at NASA are in the process of developing high-performance 7K lbf class RDRE hardware for their potential use in lander, upper stage, and even launch vehicle applications. Clear advantages have been demonstrated with AM including program cost and schedule reductions of up to 50%. It is well known that injector performance is integrally linked to the global performance of a combustion device. This is especially the case for RDREs since detonation stability is heavily dependent on the mixedness of propellants. A major goal of this work is to identify what has been done in the open experimental literature and what injectors design features are conducive to high performance in the detonation cycle. This paper reviews the available literature and reports the primary gaps in the knowledge base needed by the pressure gain combustion (PGC) community. Major conclusions are documented, and suggestions given towards the design of high-performance liquid RDRE injectors. In addition, the integration of metal AM into the design of liquid RDRE injector schemes is included.

rotating detonation rocket engine↗

NASA NTRS1972

Space Storable Propellant Performance Gas/Liquid Like-Doublet Injector Characterization

A 30-month applied research program was conducted, encompassing an analytical, design, and experimental effort to relate injector design parameters to simultaneous attainment of high performance and component (injector/thrust chamber) compatibility for gas/liquid space-storable propellants. The gas/liquid propellant combination selected for study was FLOX (82.6% F2)/ambient temperature gaseous methane. The injector pattern characterized was the like-(self)-impinging doublet. Program effort was apportioned into four basic technical tasks: injector and thrust chamber design, injector and thrust chamber fabrication, performance evaluation testing, and data evaluation and reporting. Analytical parametric combustion analyses and cold flow distribution and atomization experiments were conducted with injector segment models to support design of injector/thrust chamber combinations for hot fire evaluation. Hot fire tests were conducted to: (1) optimize performance of the injector core elements, and (2) provide design criteria for the outer zone elements so that injector/thrust chamber compatibility could be achieved with only minimal performance losses.

Falk, A. Y.↗

NASA NTRS1974

Some injector element detail effects on screech in hydrogen-oxygen rockets

An experimental investigation was conducted to learn more about how the specific details of a concentric tube injection element affect the screech characteristics of a hydrogen-oxygen rocket engine. The four variables investigated were (1) impingement angle, (2) oxidizer tube blunt base thickness, (3) oxidizer tube recess and extension, and (4) oxidizer tube-annulus concentricity. Tests were made using a 27.34-cm (10.77-in.) diameter heat sink combustor at nominally 300-psia chamber pressure. All of the test variables were investigated using a 157 element circular pattern injector. Additional oxidizer tube recess tests were made with a 421 element hexagonal pattern injector. Tests were conducted over the oxidant-fuel ratio range of 4 to 6. Stability evaluation for each configuration was made using the hydrogen temperature rating technique. Several element configurations were also cold flow tested using nitrogen and water as simulants. The element detail changes resulted in changes in hydrogen injector pressure drop even though the physical injection area was constant for all similar tests in both hot firing and cold flow tests. These changes in injector pressure drop produced changes in combustion stability. The data were correlated with a modified version of a previously reported injection area ratio correlation. By interpreting changes in injector pressure drop as changes in injector hydrogen flow resistance, the data were compared with a hydrogen flow response stability model and were found to be in agreement.

Hannum, N. P.↗

NASA NTRS1976

Liquid rocket engine injectors

The injector in a liquid rocket engine atomizes and mixes the fuel with the oxidizer to produce efficient and stable combustion that will provide the required thrust without endangering hardware durability. Injectors usually take the form of a perforated disk at the head of the rocket engine combustion chamber, and have varied from a few inches to more than a yard in diameter. This monograph treats specifically bipropellant injectors, emphasis being placed on the liquid/liquid and liquid/gas injectors that have been developed for and used in flight-proven engines. The information provided has limited application to monopropellant injectors and gas/gas propellant systems. Critical problems that may arise during injector development and the approaches that lead to successful design are discussed.

G S Gill↗

NASA NTRS1979

Driving of combustor oscillations by gaseous propellant injectors

Measurements of reactive admittances that describe the capabilities of the combustion processes associated with coaxial gaseous fuel injectors to amplify combustor oscillations are presented. These admittances are needed in (1) stability analyses of rocket motors and gas turbine combustors; (2) for the evaluation of 'driving' capabilities of injectors; and (3) for checking the applicability of a theoretical model. The modified standing-wave technique was used to determine the admittances of the combustion processes in coaxial injectors utilizing air-acetylene mixtures. The measurements indicated that (1) coaxial injectors can sustain combustion instabilities over wide frequency ranges; (2) the maximum driving capability of an injector decreases with an increased equivalence ratio; (3) the frequency at which maximum driving is observed decreases with the increased equivalence ratio; and (4) the characteristic combustion time of an injector design decreases with increased frequency.

Janardan, B. A.↗

NASA NTRS1981

A parametric study of staged fuel injector configurations for scramjet applications

A parametric study of staged (multiple) perpendicular fuel injector configurations has been conducted using a computer code which solves the two-dimensional elliptic Navier-Stokes equations. The program computes the turbulent mixing and reaction of hydrogen fuel and air and allows the study of separated regions of the flow immediately preceding and following the injectors. The validity of the code is demonstrated in a cold flow helium injection study with a single injector. Results are presented that describe the flow field near opposing staged injectors over a range of parameters. Parameters that are varied include injector size, fuel split, and distance between injectors. Comparisons of the configurations are made to assess their mixing and potential flame-holding qualities.

Weidner, E. H.↗

NASA NTRS1981

Effect of fuel injector type on performance and emissions of reverse-flow combustor

The combustion process in a reverse-flow combustor suitable for a small gas turbine engine was investigated to evaluate the effect of fuel injector type on performance and emissions. Fuel injector configurations using pressure-atomizing, spill-flow, air blast, and air-assist techniques were compared and evaluated on the basis of performance obtained in a full-scale experimental combustor operated at inlet conditions corresponding to takeoff, cruise, low power, and idle and typical of a 16:1-pressure-ratio turbine engine. Major differences in combustor performance and emissions characteristics were experienced with each injector type even though the aerodynamic configuration was common to most combustor models. Performance characteristics obtained with the various fuel injector types could not have been predicted from bench-test injector spray characteristics. The effect of the number of operating fuel injectors on performance and emissions is also presented.

Norgren, C. T.↗

NASA NTRS1982

Atomization and combustion properties of flashing injectors

Flashing injection involves expanding a fluid through an injector until a supersaturated state is reached, causing a portion of the fluid to flash to a vapor. This investigation considered the flow, atomization and spreading properties of flashing injectors flowing liquids containing dissolved gases (Jet A/air) as well as superheated liquids (Freon 11). The use of a two stage expansion process, separated by an expansion chamber, was found to be beneficial for good atomization properties of flashing injection - particularly for dissolved gas systems. Both locally homogeneous and separated flow models provided good predictions of injector flow properties. Conventional correlations for drop sizes from pressure atomized and airblast injectors were successfully modified, using the separated flow model to prescribe injector exit conditions, to correlate drop size measurements. Additional experimental results are provided for spray angle and combustion properties of sprays from flashing injectors.

Solomon, A. S. P.↗

NASA NTRS1985

Flow and atomization in flashing injectors

Flashing injection involves expanding a fluid through an injector until a supersaturated state is reached, causing a portion of the fluid to flash to a vapor. This investigation considered the flow, atomization and spreading properties of flashing injectors with flowing liquids containing dissolved gases (Jet A/air) as well as superheated liquids (Freon-11). The use of a two-stage expansion process, separated by an expansion chamber, was found to be beneficial for good atomization properties of flashing injection - particularly for dissolved gas systems. Both locally homogeneous and separated flow models provided good predictions of injector flow properties. Conventional correlations for drop sizes from pressure and airblast injectors were successfully modified, using the separated flow model to prescribe injector exit conditions, and to correlate drop size measurements. Additional experimental results are provided for spray angles of sprays from flashing injectors.

Solomon, A. S. P.↗

NASA NTRS1991

Heavy hydrocarbon main injector technology program

The Heavy Hydrocarbon Main Injector Program was an analytical, design, and test program to demonstrate an injection concept applicable to an Isolated Combustion Compartment of a full-scale, high pressure, LOX/RP-1 engine. Several injector patterns were tested in a 3.4-in. combustor. Based on these results, features of the most promising injector design were incorporated into a 5.7-in. injector which was then hot-fire tested. In turn, a preliminary design of a 5-compartment 2D combustor was based on this pattern. Also the additional subscale injector testing and analysis was performed with an emphasis on improving analytical techniques and acoustic cavity design methodology. Several of the existing 3.5-in. diameter injectors were hot-fire tested with and without acoustic cavities for spontaneous and dynamic stability characteristics.

Arbit, H. A.↗

NASA NTRS1993

Effects of compression and expansion ramp fuel injector configuration on scramjet combustion and heat transfer

A scramjet combustor with four wall-ramp injectors containing Mach-1.7 fuel jets in the base of the ramps was investigated experimentally. During the test program, two swept ramp injector designs were evaluated. One swept-ramp model had 10-deg compression-ramps and the other had 10-deg expansion cavities between flush wall ramps. The scramjet combustor model was instrumented with pressure taps and heat-flux gages. The pressure measurements indicated that both injector configurations were effective in promoting mixing and combustion. Autoignition occurred for the compression-ramp injectors, and the fuel began to burn immediately downstream of the injectors. In tests of the expansion ramps, a pilot was required to ignite the fuel, and the fuel did not burn for a distance of at least two gaps downstream of the injectors. Once initiated, combustion was rapid in this configuration. Heat transfer measurements showed that the heat flux differed greatly both across the width of the combustor and along the length of the combustor.

Stouffer, Scott D.↗

NASA NTRS1998

Injector Mixing Efficiency Experiments

Various optical diagnostic techniques such as laser induce fluorescence, Raman spectroscopy, laser Doppler velocimetry, and laser light scattering have been employed to study the flowfield downstream of a single injector element in a optically accessible rocket chamber at Penn State for a number o years. These techniques have been used with both liquid and gaseous oxygen at pressures up to 1000 psia which is the limit of the facility. The purpose of the test programs at Penn State were to develop the techniques and to study the flow field from various injector designs. To extend these studies to higher pressure and ultimately to multiple injectors require the capabilities of the Marshall Space Flight Center. These studies will extend the data base available for the various injector designs to higher pressure as well as to determine the interaction between multiple injectors. During this effort the Princeton Instruments ICCD camera was set up and checked out. The functionality of the system has been thoroughly checked and the shutter compensation time was found to be not working. The controller was returned to the manufacturer for warranty repair. The sensitivity has been measured and found to be approximately 60 counts per photon at maximum gain which agrees with the test data supplied by the manufacturer. The actual value depends on wavelength. The Princeton Instruments camera was been installed in a explosion proof tube for use with the rocket combustor. A 35 mm camera was also made ready for taking still photos inside the combustor. A fiber optic was used to transmit the laser light from an argon-ion laser to the rocket combustor for the light scattering images. This images were obtained for a LOX-hydrogen swirl coax injector. Several still photos were also obtained with the 35 mm camera for these firings.

Moser, Marlow D.↗

NASA NTRS1999

Mixing Characteristics of Coaxial Injectors at High Gas to Liquid Momentum Ratios

A study of the spray of a swirl coaxial gas-liquid injector operating at high gas to liquid momentum ratios is reported. Mixing and droplet size characteristics of the swirl injector are also compared to a shear coaxial injector, currently being used in the Space Shuttle Main Engine fuel preburner. The injectors were tested at elevated chamber pressures using water as a LOX simulant and nitrogen and helium as gaseous hydrogen simulants. The elevated chamber pressure allowed for matching of several of the preburner injector conditions including; gas to liquid momentum ratio, density ratio and Mach number. Diagnostic techniques used to characterize the spray included; strobe back-light imaging, laser sheet spray imaging, mechanical patternation, and a phase Doppler interferometry. Results thus far indicate that the radial spreading of the swirl coaxial spray is much less than was reported in previous studies of swirl injectors operating at atmospheric back-pressure. The swirl coaxial spray does, however, exhibit a smaller overall droplet size which may be interpreted as an increase in local mixing.

Strakey, P. A.↗

NASA NTRS1997

Mixing Enhancement in a Lobed Injector

An experimental investigation of the non-reactive mixing processes associated with a lobed fuel injector in a coflowing air stream is presented. The lobed fuel injector is a device which generates streamwise vorticity, producing high strain rates which can enhance the mixing of reactants while delaying ignition in a controlled manner. The lobed injectors examined in the present study consist of two corrugated plates between which a fuel surrogate, CO2, is injected into coflowing air. Acetone is seeded in the CO2 supply as a fuel marker. Comparison of two alternative lobed injector geometries is made with a straight fuel injector to determine net differences in mixing and strain fields due to streamwise vorticity generation. Planar laser-induced fluorescence (PLIF) of the seeded acetone yields two-dimensional images of the scalar concentration field at various downstream locations, from which local mixing and scalar dissipation rates are computed. It is found that the lobed injector geometry can enhance molecular mixing and create a highly strained flowfield, and that the strain rates generated by scalar energy dissipation can potentially delay ignition in a reacting flowfield.

Smith, L. L.↗