Marlow D. Moser

Effect of Chamber Backpressure on Swirl Injector Fluid Mechanics

Fluid mechanics of a liquid swirl injector element at various chamber backpressures were investigated. The center-jet swirling element was designed using typical liquid propellant rocket engine parameters, then manufactured and tested in a high-pressure, optically accessible, cold flow facility. Water was injected into a chamber pressurized with gaseous nitrogen at a constant swirl injector flow rate of 0.09 kg/s. The chamber backpressure ranged from 0.10 to 4.81 MPa. The film thickness and spray angle near the nozzle exit were measured by shadowgraphy. The film thickness was also measured within the injector upstream of the exit through a transparent nozzle tube section. Increasing the chamber backpressure for this fixed mass flow rate increased the film thickness from predicted design values. Measured discharge coefficient values increased with increasing chamber backpressure, reflecting the observed increase in internal nozzle film thickness. The spray angle decreased for increasing chamber backpressure.

Direct ultrasonic measurement of solid propellant ballistics

This article illustrates the application of an ultrasonic pulse–echo technique to determine the burning rate of a composite solid propellant as a function of pressure. An evaluation of the measurement uncertainty of the method is also presented. Unlike the more traditional strand burner techniques, where dozens of constant pressure tests are necessary, the ultrasonic technique measures the burn surface position thousands of times per second as the pressure varies. This reduces the number of tests necessary to determine the ballistic characteristics of the propellant by an order of magnitude. This work presents new methods to characterize the changing speed of sound in the propellant and quantitative estimates of the measurement uncertainty in the burning rate measurement. The results of the uncertainty analysis showed that the measurement is accurate to around 4%. The propellant samples were tested in a closed-combustion vessel, under pressurization rates of up to 15.8 MPa/s. The data obtained with the cl...

Experimental Cold Flow Characterization of a Swirl Coaxial Injector Element

A full-scale swirl coaxial injector element has been designed as part of an effort to examine liquid oxygen and liquid methane (LOX-LCH4) combustion instability phenomena under lunar ascent engine operating conditions. The scope of the LOX-LCH4 study encompassed cold flow, low pressure combustion, and high pressure combustion experimentation in an effort to establish a fast response analysis methodology for evaluating injector performance. As a baseline investigation, the spray characteristics of the injector’s central LOX post, exclusive of the exterior LCH4 annulus, were evaluated at constant ambient back pressure, across a range of steady mass flow rates. The effects of mass flow rate variance on the swirling sheet’s free cone spray angle and penetration length were assessed. Droplet velocity and diameter measurements were mapped within the spray’s primary and secondary atomization regions. The impact of throttling on the injector’s atomization quality was also surveyed.

Cold Flow Testing for Liquid Propellant Rocket Injector Scaling and Throttling

Scaling and throttling of combustion devices are important capabilities to demonstrate in development of liquid rocket engines for NASAs Space Exploration Mission. Scaling provides the ability to design new injectors and injection elements with predictable performance on the basis of test experience with existing injectors and elements, and could be a key aspect of future development programs. Throttling is the reduction of thrust with fixed designs and is a critical requirement in lunar and other planetary landing missions. A task in the Constellation University Institutes Program (CUIP) has been designed to evaluate spray characteristics when liquid propellant rocket engine injectors are scaled and throttled. The specific objectives of the present study are to characterize injection and primary atomization using cold flow simulations of the reacting sprays. These simulations can provide relevant information because the injection and primary atomization are believed to be the spray processes least affected by the propellant reaction. Cold flow studies also provide acceptable test conditions for a university environment. Three geometric scales - 1/4- scale, 1/2-scale, and full-scale - of two different injector element types - swirl coaxial and shear coaxial - will be designed, fabricated, and tested. A literature review is currently being conducted to revisit and compile the previous scaling documentation. Because it is simple to perform, throttling will also be examined in the present work by measuring primary atomization characteristics as the mass flow rate and pressure drop of the six injector element concepts are reduced, with corresponding changes in chamber backpressure. Simulants will include water and gaseous nitrogen, and an optically accessible chamber will be used for visual and laser-based diagnostics. The chamber will include curtain flow capability to repress recirculation, and additional gas injection to provide independent control of the backpressure. This paper provides a short review of the appropriate literature, as well as descriptions of plans for experimental hardware, test chamber instrumentation, diagnostics, and testing.

Experimental Pulsator Characterization for Liquid Injector Research

The Propulsion Research Center at the University of Alabama in Huntsville (PRC) has designed and built a hydro-mechanical pulsator to simulate the effects of high frequency combustion instability (HFCI) in liquid rocket engines (LRE). The response characteristics (output pressure fluctuation, amplitude, and frequency) of the pulsator were evaluated in an atmospheric test rig with filtered de-ionized water as the primary fluid. The outlet of the pulsator was connected to a swirl injector LOX post to provide some downstream flow resistance. The pulsator control variables (primary flow bypass throttle, back pressure throttle, motor drive frequency and steady state injector pressure) were systematically varied to assess the dependence of the output flow characteristics on the control variables. For each test, the average mass flow rates of the waste water, seal water leakage, and fluid delivered through the injector were measured. Dynamic pressure was measured at the pulsator exit and the mean static pressure was measured at both the injector and at the pulsator inlets. High frequency pressure measurements show a periodic pressure pulsation with maximum peak to peak amplitude of 2% of the injector pressure when the pulsator is activated. The preliminary characterization of the pulsator was insufficient to fully characterize the pulsator; however, general trends showing the relationship between the control variables are visible in the data. The relationships between the variables show that higher order curve fits are required but the amount of data collected in this initial characterization is insufficient to generate those relationships.

Ballistic Properties of Mixed Hybrid Propellants

The objective of this project is to experimentally characterize the ballistic properties of a mixed-hybrid propulsion concept that is both controllable and safe. The propulsion configuration was a center-perforated grain with gaseous oxidizer injected at the head end. The fuel grains were loaded with low levels of oxidizer and additives that enhanced their density, and increased the solid regression rate. The propellant formulations have been formulated using a design of experiments approach that evaluates the effects of three ingredients. They were evaluated at the University of Alabama in Huntsville Propulsion Test Facility to determine fundamental ballistic properties that are necessary inputs to designers for system evaluations. The results show that the ingredients added to the solid grain produces up to a 150-300% augmentation in the solid burning rate when compared to grains without any additives at the same flux and pressure levels. The burning rate was determined to be both a function of pressure and oxidizer flux. The propellant with the 25% ammonium perchlorate and the additive burned the fastest. The propellants will stop burning when the gaseous oxidizer flow is stopped.

Effects of Variable Chamber Pressure on Swirl Coaxial Injection: A Cold Flow Study

A geometrically full-scale swirl injector element was subjected to cold flow testing to explore the influence of steady chamber pressure on spray characteristics and selfatomization at conditions that reflect real liquid rocket combustor conditions. Propellant simulant flow rate was held constant at 0.82 kg/s over a chamber pressure range of 0.95 MPa to 3.45 MPa. Imaging studies observed free cone spray angle to deviate from the theoretical value when conditioned with positive variation in chamber pressure. Additionally, sheet penetration length decreased as a function of chamber pressure conditions. Phase Doppler Particle Analyzer measurements resolved droplet velocity and diameter profiles through the spray envelope. Analysis showed mean velocities and diameters in the primary atomization region to behave inversely to chamber pressure.

Pulse-echo measurements of unsteady propellant deflagration

This article discusses a laboratory test method for the measurement of nonsteady deflagration rates and combustion-stability properties of solid propellants. The method combines ultrasonic pulse-echo measurement, fluidic pressure modulation, and digital signal processing techniques to compute the real and imaginary components of pressure-coupled response functions of solid propellants. A description of the apparatus identifies the major components and their functions. A section on the experimental procedure illustrates the steps necessary to conduct a test, and a section on the data reduction technique describes the sequence of steps employed to compute the nonsteady burn rate and the pressure-coupled response. A detailed implementation of the technique as it applies to two representative tests follows, in which graphs and tables illustrate the intermediate data reduction steps. Details of the evaluation of the uncertainties associated with the technique and the data reduction algorithms are also presented.

UAH Solid Propellant Characterization

One of the main properties of solid propellants that must be known before a solid propellant design can be implemented into any use is the burning rate law. The objective of this works is to investigate the temperature sensitivity as a function of pressure for a composite propellant. An ultrasonic pulse-echo device measured the burning rate of 20-gram samples in a closed bomb at pressures from 250 to 3,000 pounds per square inch. Repeated tests show sample-to-sample variations in the burning rate measurements. The standard deviation of burning rates at a reference condition ranged from 0.4% to 0.6% at initial propellant temperatures of 145F and 75F respectively.

An Advanced Digital Cross Correlation Method for Solid Propellant Burning Rate Determination

The work presented here is a study of a digital method for determining the combustion bomb burning rate of a fuel-rich gas generator propellant sample using the ultrasonic pulse-echo technique. The advanced digital method, which places user defined limits on the search for the ultrasonic echo from the burning surface, is computationally faster than the previous cross correlation method, and is able to analyze data for this class of propellant that the previous cross correlation data reduction method could not. For the conditions investigated, the best-fit burning rate law at 800 psi from the ultrasonic technique and advanced cross correlation method is within 3 percent of an independent analysis of the same data, and is within 5 percent of the best-fit burning rate law found from parallel research of the same propellant in a motor configuration.