ALBERT

Description: Testing of a powerpack configuration (turbomachinery and gas generator assembly) and the first complete engine system of the liquid oxygen/liquid hydrogen propellant J-2X rocket engine have been completed at the NASA Stennis Space Center. The combustion stability characteristics of the gas generator assemblies on these two systems are of interest for reporting since considerable effort was expended to eliminate combustion instability during early development of the gas generator assembly with workhorse hardware. Comparing the final workhorse gas generator assembly development test data to the powerpack and engine system test data provides an opportunity to investigate how the nearly identical configurations of gas generator assemblies operate with two very different propellant supply systems one the autonomous pressure-fed test configuration on the workhorse development test stand, the other the pump-fed configurations on the powerpack and engine systems. The development of the gas generator assembly and the elimination of the combustion instability on the pressure-fed workhorse test stand have been reported extensively in the two previous Liquid Propulsion Subcommittee meetings 1-7. The powerpack and engine system testing have been conducted from mid-2011 through 2012. All tests of the powerpack and engine system gas generator systems to date have been stable. However, measureable dynamic behavior, similar to that observed on the pressure-fed test stand and reported in Ref. [6] and attributed to an injection-coupled response, has appeared in both powerpack and engine system tests. As discussed in Ref. [6], these injection-coupled responses are influenced by the interaction of the combustion chamber with a branch pipe in the hot gas duct that supplies gaseous helium to pre-spin the turbine during the start transient. This paper presents the powerpack and engine system gas generator test data, compares these data to the development test data, and provides additional combustion stability analyses of the configurations.

Description: As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center designed, fabricated, assembled and hot-fire tested an oxygen/hydrocarbon propellant multi-element integrated test article that included an oxidizer-rich oxygen/hydrocarbon propellant preburner and a staged-combustion main injector. Also as part of this project, the Air Force Research Laboratory fabricated single-element main injectors of the same designs as used in the NASA multi-element injectors, and tested them in a staged-combustion integrated test article that used an oxidizer-rich oxygen/hydrogen propellant preburner. Final results of the multi-element and single-element staged-combustion main injector test programs are described in companion papers at this JANNAF meeting. The design, development, and preliminary test results of these main injectors have also been described in previous JANNAF papers. The main injector element designs were all based on relatively conventional gas-centered swirl coaxial injector element configurations such as used in Russian RD-170 and NK-33 engines, and planned for use in future U.S.-built experimental engine systems such as the Hydrocarbon Boost program demonstration engine. Four different elements were tested in both the multi-element and single-element main injectors, at similar combustion chamber pressures, chamber contraction ratios, and mixture ratios. Variations of the element features included recess depth, fuel gap width, and the presence of the sleeve separating the swirling fuel flow from the axial oxidizer flow. This paper compares the hydraulics, combustion performance, stability, and compatibility characteristics of the single-element and multi-element injectors operated at similar conditions. The single-element hardware is shown to have captured a significant level of the operability of the multi-element hardware.

Description: General guidelines are provided in CPIA Publication 6551 for evaluation of the response due to an artificial disturbance, however the guideline also allows for ambiguous interpretation. Stability rating devices that produce an artificial disturbance are traditionally explosive bombs and pulse guns. This paper reviews a recently developed objective process that can be applied consistently in the reduction of artificial disturbance dynamic data. It also examines three methods of response evaluation. The first method examines the response of a specific mode of interest and requires data filtering encompassing that mode. The second method examines the response of a specific mode of interest and its nonlinear components and requires a more complex filtering scheme. The third method examines the response of the entire dynamic system and consists of examining a wide bandwidth consisting of multiple modes of interest. The evaluation process is described and the advantages and disadvantages of the evaluation methods are discussed. Signal processing is used as a tool in quantifying the assessment, clearly as an improvement from the subjective heritage approach consisting primarily of engineering judgement. Data for several engines and components have been compiled and evaluated using these methods. A summary of these combustion devices is provided and observations are discussed.

Description: In 2015 and 2016, the National Aeronautics and Space Administration Marshall Space Flight Center designed, fabricated, assembled and hot-fire tested an oxygen/RP-1 propellant multi-element oxidizer-rich staged-combustion test article. The main objective was to provide thrust chamber combustion stability data as part of the Combustion Stability Tool Development program, although demonstration of performance and compatibility of oxidizer-rich main injectors was also important. Funding was provided by the Air Force Space and Missile Systems Center. Five configurations of main injectors were designed and fabricated, using conventional gas-centered swirl coaxial injector element designs generally similar to those used in oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines. Variations of element features included element size, recess depth, fuel gap width, and the presence of the sleeve separating the swirling fuel flow from the axial oxidizer flow. Ablative combustion chambers were fabricated based on hardware previously used at the NASA MSFC for testing at similar size and pressure. Existing oxygen/RP-1 oxidizer-rich subscale preburner injectors and hot gas ducts from a previous NASA-funded program were modified for use to supply the oxidizer-rich combustion products to the oxidizer circuit of the main injector of the thrust chamber. Testing of the resulting integrated test article - which included the preburner, inter-connecting hot gas duct, main injector, and ablative combustion chamber - was conducted at Test Stand 116 at the East Test Area of the NASA MSFC. The test article was well instrumented with static and dynamic pressure, temperature, and vibration sensors. This paper presents and discusses all the hot-fire test results of the integrated test article thrust chamber. Eighteen successful hot-fire tests of the integrated rig were conducted. Testing was accomplished with all five of the injector element concepts. Main combustion chamber pressures ranged from 710 to 2350 psia, and main combustion chamber mixture ratios ranged from 2.47 to 2.87. A chamber barrier fuel film coolant of about 2% to 4% of the total fuel flow was used for most tests. Characteristic exhaust velocity efficiency excluding the influence of the fuel film cooling ranged from 91% to 98% of theoretical. All tests of the thrust chamber exhibited stable combustion, even down to 40% of nominal operating pressures. Compatibility of the injector face and combustion chamber walls was acceptable. This paper is a follow-on to publication of preliminary test data presented at the 2016 JANNAF Liquid Propulsion Subcommittee meeting.

Description: This work presents techniques for liquid rocket engine combustion stability assessments with respect to spontaneous stability and rough combustion. Techniques covering empirical parameter extraction, which were established in prior works, are applied for three additional programs: the F-1 Gas Generator (F1GG) component test program, the RS-84 preburner component test program, and the Marshall Integrated Test Rig (MITR) program. Stability assessment parameters from these programs are compared against prior established spontaneous stability metrics and updates are identified. Also, a procedure for comparing measured with predicted mode shapes is presented, based on an extension of the Modal Assurance Criterion (MAC).

Description: As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center was contracted to assemble and hot-fire test a multi-element integrated test article demonstrating combustion characteristics of an oxygen/hydrocarbon propellant oxidizer-rich staged-combustion engine thrust chamber. Such a test article simulates flow through the main injectors of oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines, or future U.S.-built engine systems such as the Aerojet-Rocketdyne AR-1 engine or the Hydrocarbon Boost program demonstration engine. For the thrust chamber assembly of the test article, several configurations of new main injectors, using relatively conventional gas-centered swirl coaxial injector elements, were designed and fabricated. The design and fabrication of these main injectors are described in a companion paper at this JANNAF meeting. New ablative combustion chambers were fabricated based on hardware previously used at NASA for testing at similar size and pressure. An existing oxygen/RP-1 oxidizer-rich subscale preburner injector from a previous NASA-funded program, along with existing and new inter-connecting hot gas duct hardware, were used to supply the oxidizer-rich combustion products to the oxidizer circuit of the main injector of the thrust chamber. Results from independent hot-fire tests of the preburner injector in a combustion chamber with a sonic throat are described in companion papers at this JANNAF conference. The resulting integrated test article - which includes the preburner, inter-connecting hot gas duct, main injector, and ablative combustion chamber - was assembled at Test Stand 116 at the East Test Area of the NASA Marshall Space Flight Center. The test article was well instrumented with static and dynamic pressure, temperature, and acceleration sensors to allow the collected data to be used for combustion analysis model development. Hot-fire testing was conducted with main combustion chamber pressures ranging from 1400 to 2100 psia, and main combustion chamber mixture ratios ranging from 2.4 to 2.9. Different levels of fuel film cooling injected from the injector face were examined ranging from none to about 12% of the total fuel flow. This paper presents the hot-fire test results of the integrated test article. Combustion performance, stability, thermal, and compatibility characteristics of both the preburner and the thrust chamber are described. Another companion paper at this JANNAF meeting includes additional and more detailed test data regarding the combustion dynamics and stability characteristics.

Description: The Common Extensive Cryogenic Engine program demonstrated the operation of a deep throttling engine design. The program, spanning five years from August 2005 to July 2010, funded testing through four separate engine demonstration test series. Along with successful completion of multiple objectives, a discrete response of approximately 4000 Hz was discovered and explored throughout the program. The typical low-amplitude acoustic response was evident in the chamber measurement through almost every operating condition; however, at certain off-nominal operating conditions, the response became discrete with higher amplitude. This paper summarizes the data reduction, characterization, and analysis of the 4,000 Hz response for the entire program duration, using the large amount of data collected. Upon first encountering the response, new objectives and instrumentation were incorporated in future test series to specifically collect 4,000 Hz data. The 4,000 Hz response was identified as being related to the first tangential acoustic mode by means of frequency estimation and spatial decomposition. The latter approach showed that the effective node line of the mode was aligned with the manifold propellant inlets with standing waves and quasi-standing waves present at various times. Contour maps that contain instantaneous frequency and amplitude trackings of the response were generated as a significant improvement to historical manual approaches of data reduction presentation. Signal analysis and dynamic data reduction also uncovered several other features of the response including a stable limit cycle, the progressive engagement of subsequent harmonics, the U-shaped time history, an intermittent response near the test-based neutral stability region, other acoustic modes, and indications of modulation with a separate subsynchronous response. Although no engine damage related to the acoustic mode was noted, the peak-to-peak fluctuating pressure amplitude achieved 12.1% of the mean chamber pressure at its highest. The identification of this response in terms of an instability is also discussed.

Description: The National Aeronautics and Space Administration (NASA) and Pratt & Whitney Rocketdyne are developing a liquid oxygen/liquid hydrogen rocket engine for future upper stage and trans-lunar applications. This engine, designated the J-2X, is a higher pressure, higher thrust variant of the Apollo-era J-2 engine. The contract for development was let to Pratt & Whitney Rocketdyne in 2006. Over the past several years, development of the gas generator for the J-2X engine has progressed through a variety of workhorse injector, chamber, and feed system configurations on the component test stand at the NASA Marshall Space Flight Center (MSFC). Several of the initial configurations resulted in combustion instability of the workhorse gas generator assembly at a frequency near the first longitudinal mode of the combustion chamber. In this paper, several aspects of these combustion instabilities are discussed, including injector, combustion chamber, feed system, and nozzle influences. To ensure elimination of the instabilities at the engine level, and to understand the stability margin, the gas generator system has been modeled at the NASA MSFC with two techniques, the Rocket Combustor Interaction Design and Analysis (ROCCID) code and a lumped-parameter MATLAB(TradeMark) model created as an alternative calculation to the ROCCID methodology. To correctly predict the instability characteristics of all the chamber and injector geometries and test conditions as a whole, several inputs to the submodels in ROCCID and the MATLAB(TradeMark) model were modified. Extensive sensitivity calculations were conducted to determine how to model and anchor a lumped-parameter injector response, and finite-element and acoustic analyses were conducted on several complicated combustion chamber geometries to determine how to model and anchor the chamber response. These modifications and their ramification for future stability analyses of this type are discussed.

Description: As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center was contracted to assemble and hot-fire test a multi-element integrated test article demonstrating combustion characteristics of an oxygen/hydrocarbon propellant oxidizer-rich staged-combustion engine thrust chamber. Such a test article simulates flow through the main injectors of oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines, or future U.S.-built engine systems such as the Aerojet-Rocketdyne AR-1 engine or the Hydrocarbon Boost program demonstration engine. To supply the oxidizer-rich combustion products to the main injector of the integrated test article, existing subscale preburner injectors from a previous NASA-funded oxidizer-rich staged combustion engine development program were utilized. For the integrated test article, existing and newly designed and fabricated inter-connecting hot gas duct hardware were used to supply the oxidizer-rich combustion products to the oxidizer circuit of the main injector of the thrust chamber. However, before one of the preburners was used in the integrated test article, it was first hot-fire tested at length to prove it could provide the hot exhaust gas mean temperature, thermal uniformity and combustion stability necessary to perform in the integrated test article experiment. This paper presents results from hot-fire testing of several preburner injectors in a representative combustion chamber with a sonic throat. Hydraulic, combustion performance, exhaust gas thermal uniformity, and combustion stability data are presented. Results from combustion stability modeling of these test results are described in a companion paper at this JANNAF conference, while hot-fire test results of the preburner injector in the integrated test article are described in another companion paper.

Description: Considerable interest lies in the ability to characterize the onset of spontaneous instabilities within liquid propellant rocket engine (LPRE) combustion devices. Linear techniques, such as fast Fourier transforms, various correlation parameters, and critical damping parameters, have been used at great length for over fifty years. Recently, nonlinear time series methods have been applied to deduce information pertaining to instability incipiency hidden in seemingly stochastic combustion noise. A technique commonly used in biological sciences known as the Multifractal Detrended Fluctuation Analysis has been extended to the combustion dynamics field, and is introduced here as a data analysis approach complementary to linear ones. Advancing, a modified technique is leveraged to extract artifacts of impending combustion instability that present themselves a priori growth to limit cycle amplitudes. Analysis is demonstrated on data from J-2X gas generator testing during which a distinct spontaneous instability was observed. Comparisons are made to previous work wherein the data were characterized using linear approaches. Verification of the technique is performed by examining idealized signals and comparing two separate, independently developed tools.