Hybrid rockets with mixed N2O/CO2 oxidizers for Mars Ascent Vehicles

doi:10.1016/j.actaastro.2020.05.060

Acta Astronautica

Volume 175, October 2020, Pages 254-267

https://doi.org/10.1016/j.actaastro.2020.05.060 Get rights and content

Highlights

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Mars Ascent Vehicle needs in-situ propellant production by using hybrid rockets.
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Mars has 96% carbon dioxide in atmosphere that can be used as an in-situ oxidizer compound in hybrid rockets.
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Carbon dioxide can be used as an in-situ oxidizer compound on Mars.
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Carbon dioxide has promising results on regression rate increase.
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Micron size aluminum is added to paraffin wax increases regression rate.
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Carbon dioxide can only burn metallic powders such as aluminum which is the goal of this research.

Abstract

The objective of this research is to demonstrate the feasibility of sustained combustion in hybrid rocket motor configuration with oxidizers containing high concentrations of $C O_{2}$ , an in situ propellant suitable for Mars applications. Motor tests are performed by using $N_{2} O / C O_{2}$ mixtures as the oxidizer and paraffin fuel loaded with high concentrations metal powders. Fuel grains with aluminum and magnesium powders have been cast at 40% by mass in the paraffin wax binder. $C O_{2}$ is mixed with liquid $N_{2} O$ up to 50% in weight percentage as a saturated liquid mixture. Experiments are performed in the blowdown mode using the self-pressurizing capability of the mixed oxidizers. A lab-scale hybrid motor with 70 mm grain length and 31 mm outer grain diameter is used to measure the key performance parameters such as regression rate and combustion efficiency. Successful ignition and sustained/stable combustion with the mixed oxidizer are achieved at $C O_{2}$ concentrations up to 43% by mass in the liquid mixture. The primary conclusion from this study is that the operation of hybrid rockets with substantial levels of in situ propellant $C O_{2}$ is feasible.

Introduction

Among all the planets in our solar system, Mars is the best candidate to perform human spaceflight. Mars has the most active volcanic mountains and largest impact craters among all planets. Moreover, there is methane leakage between rocks that may be an indicator of microbial life. There is also good evidence that liquid water might have flowed on the surface of Mars billions of years ago. Furthermore, Mars is relatively close to Earth compared to Saturn and Jupiter. Although Venus is another planet closer to the Earth, it has a very harsh atmosphere (high temperature, high density, corrosive …) making the planet's surface almost impossible to survive.

Two-way mission to Mars takes at least one year and requires a six-months stay on the surface. Propulsion system design is the most critical subsystem for returning to Earth after a six-month storage on the planet surface. Since it is very costly to take the propellant that will be used on the return leg all the way from Earth, in-situ propellant use is essential for the ascent vehicle that is needed for a comeback mission to Earth. There would also be a need for propulsion to transfer samples from one location to another during the long stay on the Martian surface.

Mars has a very thin atmosphere with a density of only 0.014 $k g / m^{3}$ and a pressure of merely 6 mbar. Thus, it only allows for light air vehicles such as gliders and micro helicopters [1,2]. However, these vehicles can only be used for observation purposes, and they are not expected to have the ability to transport larger payloads across the Martian surface. Thus, there is a need to have enhanced propulsion systems based on in-situ propellants that would enable practical operations such as ballistic hoppers or Mars Ascent Vehicles (MAVs). Although, both airbreathing and rocket propulsion systems can be used in air transportations systems on Mars, airbreathing engine concepts needs very large inlet areas due to the low atmospheric pressure of the planet. In addition, a high-speed vehicle is needed for a ramjet operation, since the condensed phase combustion products makes the turbojet engine impractical. Therefore, rocket propulsion systems come into forefront as a practical candidate for Mars applications.

An increased attention has been given to extract natural resources of Mars for rocket propellant production. One option is to use $C O_{2}$ from the Martian atmosphere as an oxidizer compound for propulsion systems. Although carbon dioxide acts as a fire extinguisher agent for hydrocarbon-based fuels, it can be burned with metallic fuels resulting high levels of combustion energy. Therefore, heavy loadings of metallic additives such as aluminum or magnesium is required in liquid or solid hydrocarbon fuels.

Advanced rocket propulsion concepts using $C O_{2}$ as oxidizer, and metals as the primary fuel have been studied by many researches. Boiron [3] emphasized $C O_{2}$ decomposition technique by using chemical reactors operated on the surface of the Mars for Mars Sample Return (MSR) missions. In addition, Shafirovich and Goldshleger [4,5] studied thermodynamic models of pure metallic fuel combustion such as lithium ( $L i)$ , beryllium $(B e)$ , boron $(B)$ , magnesium $(M g)$ , aluminum $(A l$ ) along with several metal hydrides such as beryllium hydride $(B e H_{2})$ , magnesium hydride $(M g H_{2})$ , diborane $(B_{2} H_{6})$ , and silane $(S i H_{4})$ . Shafirovich concluded that, despite its low specific impulse performance, magnesium is the most promising practical candidate due to its easy ignitability with $C O_{2}$ at low temperatures around 1000K.

Robert Zubrin performed theoretical performance calculations with liquid hydrides such as diborane and silane [3]. Primary challenge of these propellant combinations is their high mass fraction of condensed phase products and low $I_{s p}$ performance. According to Zubrin, $S i H_{4} / C O_{2}$ reaction provides lower condensed phase products leading to a notable increase in the $I_{s p}$ values. Although $S i H_{4}$ is storable and stable, its toxic and there is not enough information on the combustion with $C O_{2}$ as stated by Boiron [3].

An experimental study on combustion characteristics of carbon dioxide with magnesium is studied by Yue Lie [6] for rocket engines. He studied multiple gas injection mechanisms into the combustion chamber at high pressures. Yue presented the thermodynamic calculations for the combustion process of the multiphase flow environment in a lab scale rocket engine used in the experiments.

Mars Ascent Vehicle design concepts are also studied by many researchers. Carter [7] discussed technology requirements for propulsion systems such as displacement pumps and bladder lined composite tanks. One of the experimental works for a potential MAV is presented by Karp with paraffin-based fuel and Mixed Oxides of Nitrogen (MON-3) oxidizer [8]. Furthermore, Evans and Karabeyoglu also studied MON based oxidizer for MAV experiment with metallized SP7 fuel by using 30 μm sized aluminum powder [9]. SP7 paraffin-based solid fuel is developed by Space Propulsion Group, Inc., specifically for this program.

A new concept was proposed by Karabeyoglu as a hybrid rocket system using nitrous oxide and carbon dioxide mixture as the oxidizer. High concentration of the in situ component $C O_{2}$ would significantly reduce the mass needed to be brought to Mars for a specific payload. Since $N_{2} O$ and $C O_{2}$ have similar physical properties, they form ideal mixtures with fluidic properties very much like the pure components. The main idea is to use a very high metal content in paraffin-based fuels to achieve the combustion reaction with $C O_{2}$ . During the combustion, paraffin would react with $N_{2} O$ and melt the metal oxide layers so that the $C O_{2}$ component in the oxidizer can react freely with the metallic fuel [10].

Compared to solid and liquid rocket propulsion systems, hybrid propulsion system offers substantial power and mass savings and compatibility at low temperatures of Mars. In addition, throttling, shut down and re-ignitability of hybrid systems introduces mission flexibility with non-hazardous manufacturing. Operation at low temperature capability, long oxidizer storage without risk of decreased performance are key factors that makes hybrid ideal candidates for in-space missions [11].

Hybrid rockets commonly use HTPB (Hydroxyl-Terminated Polybutadiene) and paraffin-based fuels that are widely available, cost effective and easy to handle. Paraffin-based fuels show 3 to 5 times higher regression rate at similar mass fluxes compared to HTPB and other polymeric fuels due to the additional entrainment mass transfer mechanism at the burning surface [12]. When the paraffin-based fuel burns, it produces a liquid layer on the fuel surface which has low viscosity and low surface tension. This liquid layer becomes hydrodynamically unstable by the gas flow in the fuel port and this instability leads to the lift-off of fuel droplets from the surface increasing the overall mass transfer rate. Moreover, the hydrophobic nature of paraffin wax makes it an ideal binder for metals and metal hydrides such as aluminum hydride. Easy and safe addition of metals help achieving higher specific impulse performance and fuel densities with paraffin-based fuels. In addition, low glass transition temperature of paraffin wax (around −180 °C) makes it a better binder for the cold Mars environment.

$N_{2} O_{4}$ , $H_{2} O_{2}$ , liquid oxygen $(L O X)$ are common oxidizers that are used in hybrid rocket propulsion. They have certain shortcomings compared to nitrous oxide, which is another oxidizer extensively used in this field. Specifically, nitrous oxide is non-toxic, non-corrosive and easy to handle compared to $N_{2} O_{4}$ and $H_{2} O_{2}$ , and unlike $L O X$ , $N_{2} O$ is highly storable and can be loaded well before the launch operation. Nitrous oxide is a self-pressurizing agent that eliminates the need of complex and costly pressurization systems or turbopumps. In addition, saturation properties of $C O_{2}$ including density, vapor pressure, and viscosity are very similar to $N_{2} O$ . Advantages of $C O_{2} / N_{2} O$ mixture is presented by Karabeyoglu [10] as (i) improved $I_{s p}$ performance compared to pure $C O_{2}$ , (ii) decreased two-phase losses due to reduced mass fraction of condensed phase products, (iii) low freezing point of the oxidizer mixture is ideal for Martian environment (iv) less complicated and lighter compared to liquid bipropellant engines. Considering all presented factors, the mixture of $C O_{2} a n d N_{2} O$ is a promising practical oxidizer option for a Mars system utilizing in situ propellants. However, for this concept to work, there is a need to add metallic powders to the fuel grain at very concentration levels, since the carbon dioxide can only combust with metals.

The goal of this paper is to study the feasibility of heavily metal loaded hybrid rocket motors operating with the $C O_{2} / N_{2} O$ oxidizer mixture by conducting motor tests using a lab-scale hybrid system. A large fraction of the tests that has been performed in this research utilized aluminum powders with 3-μm average diameter (spherical shape) loaded at 40% by mass in paraffin wax. The effect of using magnesium powder is also studied in order to establish a performance comparison to the aluminum loaded fuels.

Section snippets

Metallic powder additives comparison

In rocket systems, metallic powders serve as excellent fuel additives due to their significant volumetric and gravimetric heat release during the combustion process [13]. Purity, size and shape of metallic powders directly affect the combustion performance as they control the ignition delay and the formation of the condensed combustion product (CCP). Among all other metallic powder additives, Boron is a common additive that has high combustion energy per unit mass/volume. However, it is shown

Paraffin wax properties

Paraffin wax is selected as main binder for fuel formulations. Benefits of paraffin waxes are as follows (1) Cost effective and readily available in the market, (2) Hydrophobic: Protects metal additives from the water vapor in air. (3) High performance due to high hydrogen/carbon ratio. (4) Three to five times higher regression rate than classical polymeric fuels due to liquid droplet entertainment mechanism. Therefore, faster regression rate leads to the circular port grain designs which leads

Oxidizer mixing process

Although $N_{2} O$ and $C O_{2}$ have similar fluidic characteristics, mixing of two self-pressurizing liquid oxidizers (saturated liquids) requires a careful process. Since the initial oxidizer mass during the experiments is around 210 g, addition of 10% $C O_{2}$ (21 g) requires tight control for accurate results.

The oxidizer mixing process is started with filling the $N_{2} O$ into the run tank at a level higher than the desired quantity. When the tank is fully filled, tank is vented in order to obtain the required

Experimental setup

Test setup consists of a hybrid rocket motor with a stainless-steel case that is 122 mm long and has a 44.5 mm outside diameter. The pre combustion chamber is 20 mm long. Combustion chamber consist of 70 mm long fuel grain. The cross-sectional view of the hybrid motor is presented in Fig. 2.

The motor is ignited using a pyrogen which itself is ignited with a heated nichrome wire. Pyro consists of a 3-g KN based solid charge. Oxidizer flow rate in the hybrid rocket motor varies between 15 and

Hybrid motor performance evaluation

In order to determine the oxidizer flow rate for the $N_{2} O / C O_{2}$ oxidizer mixture, Eq. (6) is used with injector discharge coefficient ( $c_{d}$ ), cross sectional area of the individual injector hole ( $A_{i n j}$ ), number of injector holes $(N_{i n j})$ , average fluid density at the injector ( ${\overline{ρ}}_{l i q u i d}$ ), the difference of average chamber pressure $({\overline{P}}_{c h a m b e r})$ and average tank pressure $({\overline{P}}_{t a n k})$ . Pressure drop from the tank to the injector face is believed to be small for the test set up and this effect is ignored in the

Test data relative error results

An error analysis has been conducted in order to understand the uncertainties in the performance parameters of the hybrid rocket motor. The relative errors can be found by defining nondimensional parameters at the thrust termination. The goal is to determine the actual regression rate during the full pressure burn by eliminating the effect of regression rate that took place during the thrust termination phase of combustion. The error analysis used in this research is based on Karabeyoglu [22,23

Test codes

Experiments are performed with various fuel combinations that have specific codes. Table 4 illustrates Propellant Characterization Test (PCT) code used for the test data. The fuel code starts with the metallic powder name, metallic powder percentage by mass, structural additive percentage added in the paraffin wax to increase mechanical properties and number of fuels produced. In addition, spin cast or axial cast grain is also remarked in the fuel code.

PCT test results

The effect of $C O_{2}$ concentration used in

Discussion: Mars ascent vehicle propulsion system

The hybrid rockets based on fuels with large metal loadings and the $N_{2} O$ and $C O_{2}$ mixture as the oxidizer is a promising candidate for propulsion systems that can be used in Mars missions. A large fraction of the propellants can easily be pumped from the Martian atmosphere. Therefore, usage of carbon dioxide as an in situ oxidizer reduces the mass that needs to be transported from the Earth. It is important to note that the oxidizer mixture and the paraffin-based binder are highly storable under

Declaration of competing interest

The authors declare that they have no known competing financial interests or personal relationships that could have appeared to influence the work reported in this paper.

Acknowledgements

Research presented in this paper has been performed at the facilities of DeltaV Space Technologies Inc., in Istanbul Turkey. We would like to offer our special thanks to DeltaV technicians and engineers for their contributions to the project.

Ozan Kara is PhD Candidate at Mechanical Engineering Department of KOC University. His major research field is hybrid rocket motor design by using metallic additives. He is also working on producing in-situ propellant technologies on the Mars and Mars Ascent Vehicle system optimization. Mr. Kara is the Middle East Regional Coordinator at the Space Generation Advisory Council. In addition, Mr. Kara is member of several technical committees in American Institute of Aeronautics and Astronautics

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According to data obtained by Mars probes [4], the surface atmosphere of Mars is 95.3% CO2 and its polar regions have long been covered by dry ice, so collecting CO2 from the Martian surface as an oxidant can greatly reduce the payload from Earth and save exploration costs. At present, the known in situ resource utilization methods include N2O/CO2 rocket engines [5], Li/CO2 electric propulsion systems [6], and Mg/CO2 powder rocket engines [7,8]. There has been much experimental research on Mg/CO2 powder rocket engines [9–12], but there has been relatively little research on their thermal protection.
To solve the problem of the thermal protection of Mg/CO₂ powder rocket engines, how different working conditions in the regenerative-cooling channel influence the pressure drop of CO₂ coolant is investigated. It is found that the proportion of two-phase acceleration pressure drop increases with increasing CO₂ mass flux, inlet temperature, and heat flux and decreases with increasing backpressure. The ratio of acceleration pressure drop to total pressure drop is ∼0.3. With increasing mass flux, incoming flow temperature, and heat flux, the two-phase frictional pressure drop increases, and the increase of backpressure decreases the two-phase frictional pressure gradient. The factor $f_{l p}$ is introduced to correct for the local pressure drop of the experimental system, and $f_{l p}$ = 1.9819 when the two-phase coolant Reynolds number is between 20 000 and 50 000. For the unidirectionally heated cooling channel, the coolant is heated unevenly therein, so the heat-transfer mode has an important influence on the frictional pressure gradient of the coolant. It is found that $μ_{l}$ can be used to predict the frictional pressure drop of $μ_{t p}$ , but the deviation is still more than 20%. Based on experimental data, a new correlation for predicting the two-phase frictional pressure gradient is proposed. The relative deviation of the predicted data is only −7.02%, and 96.12% of the predicted data fall within the ±15% deviation band, so the prediction accuracy is greatly improved. The present research provides a theoretical basis for designing a regenerative-cooling scheme for an Mg/CO₂ powder rocket engine.
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This unique structure and operation process make hybrid rocket motor possess several advantages, including safety, simple structure, relatively low cost, simplified shutdown and restarting, large range thrust regulation, and long-time working [1–5]. Because of above characteristics, hybrid rocket motor shows good value in large launch boosters, upper stage propulsion, attitude and orbit control motors, and thrust modulation [6–9]. However, the characteristic of diffuse combustion in the hybrid rocket motor leads to the low combustion efficiency [10–12].
This paper studies the effects of liquid oxidizer flow distribution on combustion characteristics and performance of hybrid rocket motor with axial injection through numerical simulations of two-phase flow. Axial injection is widely used in hybrid rocket motor, however, the regression of burning surface may change combustion efficiency of the motor at different moments. Consequently, it is necessary for hybrid rocket motor to design injector patterns that could increase the average combustion efficiency in the whole operating process. The orifice injector plate is divided into three areas, which are recorded as zone I, zone II and zone III, respectively. Oxidizer flow distribution is achieved through changing the diameter of injector holes in various zones, and three-dimensional steady numerical simulations of two-phase flow of hybrid rocket motor with axial injector patterns at different moments are conducted. The motor adopts 98% hydrogen peroxide and hydroxyl-terminated polybutadiene as the propellants. Numerical analysis reveals that oxidizer flow distribution mainly influences the flow field structure in the pre-combustion chamber and front part of fuel grain port. Increasing oxidizer mass flow rate in zone III is beneficial for the chamber head thermal environment. Relative position in the radial direction of fuel grain inner surface and oxidizer distributed in the injector plate affects fuel regression rates, and the influence mainly exists in the front part of solid fuel grain. When the oxidizer mass flow rate in zone I is equal, the average fuel regression rate rises with the increase of oxidizer flow rate in zone II at initial moment, and decreases as the oxidizer in zone II increases at middle moment. The average combustion efficiency increases with the increase of oxidizer mass flow rate in zone II when the oxidizer in zone I remains unchanged. Both the average combustion efficiency and the chamber head thermal protection need to be considered in the design of axial injector patterns. Research results of this paper provide an important reference for the design of high performance injector patterns of hybrid rocket motor.
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Hakki Karakas is a PhD Student at Istanbul Technical University. He received his MS degree from the Department of Aeronautics and Astronautics at Stanford University and BS in Aeronautics and Astronautics Faculty from İstanbul Technical University. His research interests are developing novel hybrid rocket solid fuels, studying energetic additives for hybrid rockets and internal and external aerodynamics of rockets.

Dr. Karabeyoglu is currently a full-time faculty at the Mechanical Engineering Department of KOC University. He is the co-founder of Space Propulsion Group, Inc, a company specializing in the development of advanced hybrid rocket technologies. Dr. Karabeyoglu has performed extensive research and development activities in the field of chemical propulsion and green energy with emphasis on the creation of new concepts and their implementation to real life systems. In addition, he has started a new course on Advanced Rocket Propulsion at Stanford University. Dr. Karabeyoglu is an Associate Fellow of the American Institute of Aeronautics and Astronautics.

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Hybrid rockets with mixed N2O/CO2 oxidizers for Mars Ascent Vehicles

Highlights

Abstract

Introduction

Section snippets

Metallic powder additives comparison

Paraffin wax properties

Oxidizer mixing process

Experimental setup

Hybrid motor performance evaluation

Test data relative error results

Test codes

PCT test results

Discussion: Mars ascent vehicle propulsion system

Declaration of competing interest

Acknowledgements

Acta Astronaut.

Acta Astronaut.

Acta Astronaut.

Acta Astronaut.

Evaluation of aerodynamic performance of Mars airplane in scientific balloon experiment

Fluid Mech. Res. Int.

Aerodynamic design of a martian micro air vehicle

The concept of a rocket engine using CO2/metal propellant for Mars sample return missions

Technology development and design of a hybrid Mars ascent vehicle concept

Development and testing of SP7 fuel for Mars ascent vehicle application

Mixtures of nitrous oxide, carbon dioxide and oxygen as oxidizers for Mars applications

Hybrid rockets with mixed $N_{2} O / C O_{2}$ oxidizers for Mars Ascent Vehicles