Hybrid rocket engines are promising propulsion systems which present appealing features such as safety, low cost, and environmental friendliness. On the other hand, certain issues hamper the development hoped for. The present paper discusses approaches addressing improvements to one of the most important among these issues: low fuel regression rate. To highlight the consequence of such an issue and to better understand the concepts proposed, fundamentals are summarized. Two approaches are presented (multiport grain and high mixture ratio) which aim at reducing negative effects without enhancing regression rate. Furthermore, fuel material changes and nonconventional geometries of grain and/or injector are presented as methods to increase fuel regression rate. Although most of these approaches are still at the laboratory or concept scale, many of them are promising.

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Approaches to Low Fuel Regression Rate in Hybrid Rocket Engines

http://thermophysics.ru/pdf_doc/AutoPlay/Docs/CollectionOfManuscripts/ECTP2005paper214.pdf

Analysis of the temperature distribution in a guarded hot plate apparatus for measuring thermal conductivity

The low regression rates of classic hybrid rocket fuels lead to large internal ports that limit potential applications. This experimental study investigated the increase in regression rate that results from adding a solid oxidizer and a catalyst to a hybrid fuel grain. The configuration is named a "mixed hybrid" hybrid to signify solid oxidizer and catalyst in the grain. A design of experiments approach guided fuel formulation to systematically control levels of ammonium perehlorate from 25% to 30%, ferric oxide from 0 to 5%, and hydroxyl-terminated polybutadiene from 70% to 75%. The 1.5-in. diam. port, 12-in. long center perforated grains were burned with gaseous oxygen at pressure levels from 150 to 550 psig and port flux levels from 0.1 to 0.4 lbm/s-in. 2 . The results show that the mixed hybrid propellants burn as a function of both pressure and mass flux. A grain formulation having 27.5 % ammonium perchlorate and 2.5% ferric oxide provided the maximum burning rate augmentation (447%) among the formulations tested.

Regression rates study of mixed hybrid propellants

The spray and breakup characteristics of a swirling liquid sheet were investigated by measuring the spray angle and breakup length as the axial Weber number We l was increased up to 1554 and the ambient gas pressure up to 4.0 MPa. As the We l and ambient gas density p increased, the disturbances on the annular liquid sheet surface were amplified by the increase of the aerodynamic forces, and thus the liquid sheet disintegrated from the injector exit. The measured spray angles according to the ambient gas density differed before and after the sheet broke up. Before the liquid sheet broke up, the spray angle was almost constant; however, once the liquid sheet started to break up, the spray angle decreased. As the ambient gas density and We l increased, the increasing aerodynamic force caused the breakup length to decrease. Finally, the measured breakup lengths according to the ambient gas density and We l were compared with the results of the linear instability theory. Considering the attenuation of sheet thickness in the linear instability theory, the corrected breakup length relation agreed well with our experimental results.

Effect of ambient gas density on spray characteristics of swirling liquid sheets

Heat transfer—A review of 2004 literature

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.

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Effect of Chamber Backpressure on Swirl Injector Fluid Mechanics

Scaling and throttling of combustion devices are important capabilities to demonstrate in development of liquid rocket engines for NASA's 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.

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Cold Flow Testing for Liquid Propellant Rocket Injector Scaling and Throttling

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.

Marlow D. Moser

Papers

Regression rates study of mixed hybrid propellants

Effect of Chamber Backpressure on Swirl Injector Fluid Mechanics

Cold Flow Testing for Liquid Propellant Rocket Injector Scaling and Throttling

Experimental Cold Flow Characterization of a Swirl Coaxial Injector Element

Combustion Effects of C Soot in Ammonium Nitrate Propellants