Academic literature on the topic 'Hybrid propellant rockets – Testing'

This paper presents a view on how Dr. Wernher Von Braun laid the basis for realistic modelling and testing liquid-propellants rockets, by his PhD Thesis – a secret document in 1934, which remained classified until 1960. Understanding that better mathematical modelling is needed if these rockets are to become spaceflight vehicles, he clarified in his thesis essential issues like: maximum achievable rocket speed; Laval nozzle thrust gain; polytropic processes in the combustion chamber and nozzle; influence of equilibrium and dissociation reactions; original measurement systems for rockets test stand; engineering solutions adequate for series production of the combustion chamber – reactive nozzle assembly. The thesis provided a theoretical and experimental basis for a new concept of the rocket, having a lightweight structure; low tanks pressure; high-pressure pumps and injectors; low start speed; rocket stabilization by gyroscopic means or by active jet controls; longer engine burning time; higher jet speed. Numerous tests made even with a fully assembled rocket (the “Aggregate-I”), improved mathematical model accuracy (e.g., the maximum achievable altitude predicted for the “Aggregate-II” rocket was confirmed later in-flight tests).

Experimental sounding rockets are important contributors to aerospace engineering research. However, experimental-sounding rockets are rarely used for student research projects by institutes in India. The unavailability of rocket motors, which require complex machining and explosive propellants, is a major barrier to the use of sounding rockets in student research projects. We ran into this problem while developing a sounding rocket motor for project and learning purposes. The project focuses on designing and constructing a solid rocket motor that researchers can use as the primary propulsion unit in experimental sounding rockets. Initially, basic designs were evaluated, as various concepts of observations of propellant configuration. The accessibility and ease of use of manufacturing and casting of propellants played a significant role in determining the best propellant based on these findings, the theoretical values for combustion chamber parameters were obtained. Also, materials were chosen accordingly, and a fundamental small-scale experimental design was built and extensively tested. This small-scale motor was created by combining all of the analysis and theoretical data.At experimental testing, we got to know the thrust generated is 763.47N and the motor runs for 4.1 sec, the total mass of the propellant is maxed at 1500g which gives us the max mass flow rate of 0.65Kg/sec this is the output for our solid rocket motor.

This research focuses on theoretically and experimentally evaluating the performance of a metallized hybrid-propellant, water-breathing ramjet. The aluminum and/or magnesium particles added to a polymeric (polyester) fuel grain are hydro-reactive, using the surrounding water as an oxidizer, in addition to a source of gas. Theoretically, the metal additives significantly increase the specific impulse of the motor, and as the percentage of the hydro-reactive ingredient increases, the theoretical performance increases as well. Additionally, aluminum is more energetic than magnesium. However, it was experimentally discovered that the addition of aluminum beyond 20% resulted in a slag formation and did not increase the specific impulse. Adding 30% of magnesium was relatively favorable to aluminum due to its better reactivity, enabling the achievement of an actual specific impulse of up to 485 s at standard conditions, approximately double the performance of common solid rockets.

Propellants are energetic materials abundantly used to generate the propulsion of rockets, projectiles, or other objects. A wide range of stress-state and strain-rate has to be considered in view of predicting the mechanical behaviour of this material over its different life cycles. Propellant materials are usually studied through the use of propellant representative materials (PRM) when studied in a university laboratory. In the present work an experimental device was developed to investigate the dynamic tensile response of a PRM material. This device is based on the use of a pendulum that is employed to dynamically load the PRM sample on one end. The sample is attached to an instrumented Hopkinson bar on the other end. The data processing of Hopkinson bar point measurements combined to DIC (digital-imagecorrelation) measurements allows the stress and strain levels in the sample to be characterised. Finally these experimental results can be used to enhance the constitutive modelling of PRM materials.

The AQUASONIC project is aimed to develop a sounding rocket including a hybrid propulsion system based on the propellant combination nitrous oxide and polyethylene. It takes place in the frame of the STERN (Student Experimental Rockets) programme founded by the German Space Agency (DLR) in order to promote students in the area of launch vehicles. Main element of the project is the AQUASONIC rocket, which shall reach a flight altitude of 5-6 km and a velocity of MACH 1. All major activities like design, manufacturing, verification and, finally, the launch campaign will be performed by students. The rocket shall be launched at Esrange Space Centre (Sweden) in 2016. Thus, students are able to apply their skills and knowledge to a real project like it is conducted by the space industry or research organisations.

AbstractHybrid rocket motors have several attracting characteristics such as simplicity, low cost, safety, reliability, environmental friendliness. In particular, hybrid rockets can provide complex and flexible thrust profiles not possible with solid rockets in a simpler way than liquid rockets, controlling only a single fluid. Unfortunately, the drawback of this feature is that the mixture ratio cannot be directly controlled but depends on the specific regression rate law. Therefore, in the general case the mixture ratio changes with time and with throttling. Thrust could also change with time for a fixed oxidizer flow. Moreover, propellant residuals are generated by the mixture ratio shift if the throttling profile is not known in advance. The penalties incurred could be more or less significant depending on the mission profile and requirements. In this paper, some proposed ways to mitigate or eliminate these issues are recalled, quantitatively analysed and compared with the standard case. In particular, the addition of energetic additives to influence the regression rate law, the injection of oxidizer in the post-chamber and the altering-intensity swirling-oxidizer-flow injection are discussed. The first option exploits the pressure dependency of the fuel regression to mitigate the shift during throttling. The other two techniques can control both the mixture ratio and thrust, at least in a certain range, at the expense of an increase of the architecture complexity. Moreover, some other options like pulse width modulation or multi-chamber configuration are also presented. Finally, a review of the techniques to achieve high throttling ratios keeping motor stability and efficiency is also discussed.

High vapor pressure propellants such as nitrous oxide are widely used in experimental hybrid and liquid rockets as they can be used in a self-pressurization mode, eliminating the need for external pressurization or pumps and simplifying the design of the rocket system. This approach causes the two-phase flow in the feed system and the injector orifices, which cannot be easily modeled and accounted for in the design. A dedicated test stand has been developed to better understand how the two-phase flow of the self-pressurizing propellant impacts the mass flow characteristics, enabling the simulation of the operating conditions in the rocket engine. The injectors have been studied in the range of ΔP. The flow regimes have been identified, which can be predicted by the SPI and HEM models. It has been shown that the two-phase flow quality upstream of the injector may impact the discharge coefficient in the SPI region and the accuracy of the HEM model. It has been found that the transition to the critical flow region depends on the L/D ratio of the injector orifice. A series of conclusions can be drawn from this work to design the rocket injector with a self-pressurizing propellant to better predict the mass flow rate and ensure stable combustion.

This paper presents the development of indigenous hybrid rocket technology, using 98% hydrogen peroxide as an oxidizer. Consecutive steps are presented, which started with interest in hydrogen peroxide and the development of technology to obtain High Test Peroxide, finally allowing concentrations of up to 99.99% to be obtained in-house. Hydrogen peroxide of 98% concentration (mass-wise) was selected as the workhorse for further space propulsion and space transportation developments. Over the course nearly 10 years of the technology’s evolution, the Lukasiewicz Research Network—Institute of Aviation completed hundreds of subscale hybrid rocket motor and component tests. In 2017, the Institute presented the first vehicle in the world to have demonstrated in-flight utilization for 98% hydrogen peroxide. This was achieved by the ILR-33 AMBER suborbital rocket, which utilizes a hybrid rocket propulsion as the main stage. Since then, three successful consecutive flights of the vehicle have been performed, and flights to the Von Karman Line are planned. The hybrid rocket technology developments are described. Advances in hybrid fuel technology are shown, including the testing of fuel grains. Theoretical studies and sizing of hybrid propulsion systems for spacecraft, sounding rockets and small launch vehicles have been performed, and planned further developments are discussed.

The purpose of this paper is to present some aspects regarding the computational model and technical solutions for multistage suborbital launcher for testing (SLT) used to test spatial equipment and scientific measurements. The computational model consists in numerical simulation of SLT evolution for different start conditions. The launcher model presented will be with six degrees of freedom (6DOF) and variable mass. The results analysed will be the flight parameters and ballistic performances. The discussions area will focus around the technical possibility to realize a small multi-stage launcher, by recycling military rocket motors. From technical point of view, the paper is focused on national project “Suborbital Launcher for Testing” (SLT), which is based on hybrid propulsion and control systems, obtained through an original design. Therefore, while classical suborbital sounding rockets are unguided and they use as propulsion solid fuel motor having an uncontrolled ballistic flight, SLT project is introducing a different approach, by proposing the creation of a guided suborbital launcher, which is basically a satellite launcher at a smaller scale, containing its main subsystems. This is why the project itself can be considered an intermediary step in the development of a wider range of launching systems based on hybrid propulsion technology, which may have a major impact in the future European launchers programs. SLT project, as it is shown in the title, has two major objectives: first, a short term objective, which consists in obtaining a suborbital launching system which will be able to go into service in a predictable period of time, and a long term objective that consists in the development and testing of some unconventional sub-systems which will be integrated later in the satellite launcher as a part of the European space program. This is why the technical content of the project must be carried out beyond the range of the existing suborbital vehicle programs towards the current technological necessities in the space field, especially the European one.

Automatic altitude compensation has been a holy grail of rocket propulsion for decades. Current state-of-the-art bell nozzles see large performance decreases at low altitudes, limiting rocket designs, shrinking payloads, and overall increasing costs. Aerospike nozzles are an old idea from the 1960’s that provide superior altitude-compensating performance and enhanced performance in vacuum, but have survivability issues that have stopped their application in satellite propulsion systems. A growing need for CubeSat propulsion systems provides the impetus to study aerospike nozzles in this application. This study built two aerospike nozzles using modern 3D metal printing techniques to test aerospikes at a size small enough to be potentially used on a CubeSat. Results indicated promising in-space performance, but further testing to determine thermal limits is deemed necessary.

This work describes the development of a hybrid rocket propulsion system for a reusable sounding rocket, as part of the first phase of the UKZN Phoenix Hybrid Sounding Rocket Programme. The programme objective is to produce a series of low-to-medium altitude sounding rockets to cater for the needs of the African scientific community and local universities, starting with the 10 km apogee Phoenix-1A vehicle. In particular, this dissertation details the development of the Hybrid Rocket Performance Code (HRPC) together with the design, manufacture and testing of Phoenix-1A’s propulsion system. The Phoenix-1A hybrid propulsion system, generally referred to as the hybrid rocket motor (HRM), utilises SASOL 0907 paraffin wax and nitrous oxide as the solid fuel and liquid oxidiser, respectively. The HRPC software tool is based upon a one-dimensional, unsteady flow mathematical model, and is capable of analysing the combustion of a number of propellant combinations to predict overall hybrid rocket motor performance. The code is based on a two-phase (liquid oxidiser and solid fuel) numerical solution and was programmed in MATLAB. HRPC links with the NASA-CEA equilibrium chemistry programme to determine the thermodynamic properties of the combustion products necessary for solving the governing ordinary differential equations, which are derived from first principle gas dynamics. The combustion modelling is coupled to a nitrous oxide tank pressurization and blowdown model obtained from literature to provide a realistic decay in motor performance with burn time. HRPC has been validated against experimental data obtained during hot-fire testing of a laboratory-scale hybrid rocket motor, in addition to predictions made by reported performance modelling data. Development of the Phoenix-1A propulsion system consisted of the manufacture of the solid fuel grain and incorporated finite element and computational fluid dynamics analyses of various components of the system. A novel casting method for the fabrication of the system’s cylindrical single-port paraffin fuel grain is described. Detailed finite element analyses were performed on the combustion chamber casing, injector bulkhead and nozzle retainer to verify structural integrity under worst case loading conditions. In addition, thermal and pressure loading distributions on the motor’s nozzle and its subsequent response were estimated by conducting fluid-structure interaction analyses. A targeted total impulse of 75 kNs for the Phoenix-1A motor was obtained through iterative implementation of the HRPC application. This yielded an optimised propulsion system configuration and motor thrust curve.
Thesis (M.Sc.Eng.)-University of KwaZulu-Natal, Durban, 2013.

The University of KwaZulu-Natal (UKZN) Phoenix Hybrid Sounding Rocket Programme was established in 2010. The programme’s main objective is to develop a sounding rocket launch capability for the African scientific community, which currently lacks the ability to fly research payloads to the upper atmosphere. In this dissertation, UKZN’s in-house Hybrid Rocket Performance Simulator (HYROPS) software is used to improve the design of the Phoenix-2A vehicle, which is intended to deliver a 5 kg instrumentation payload to an apogee altitude of 100 km. As a benchmarking exercise, HYROPS was first validated by modelling the performance of existing sub-orbital sounding rockets similar in apogee to Phoenix-2A. The software was found to approximate the performance of the published flight data within 10%. A generic methodology was then proposed for applying HYROPS to the design of hybrid propellant sounding rockets. An initial vehicle configuration was developed and formed the base design on which parametric trade studies were conducted. The performance sensitivity for varying propulsion and aerodynamic parameters was investigated. The selection of parameters was based on improving performance, minimising cost, safety and ease of manufacturability. The purpose of these simulations was to form a foundation for the development of the Phoenix-2A vehicle as well as other large-scale hybrid rockets. Design chamber pressure, oxidiser-to-fuel ratio, nozzle design altitude, and fin geometry were some of the parameters investigated. The change in the rocket’s propellant mass fraction was the parameter which was found to have the largest effect on performance. The fin and oxidiser tank geometries were designed to avoid fin flutter and buckling respectively. The oxidiser mass flux was kept below 650 kg/m2s and the pressure drop across the injector relative to the chamber pressure was maintained above 15% to mitigate the presence of combustion instability. The trade studies resulted in an improved design of the Phoenix-2A rocket. The propellant mass of the final vehicle was 30 kg less than the initial conceptual design and the overall mass was reduced by 25 kg. The Phoenix-2A vehicle was 12 m in length with a total mass of 1006 kg. The fuel grain length of Phoenix-2A was 1.27 m which is approximately 3 times that of Phoenix-1A. The benefit of aluminised paraffin wax as a fuel was also investigated. The results indicated that more inert mass can be delivered to the target apogee of 100 km when using a 40% aluminised paraffin wax.
M.Sc.Eng. University of KwaZulu-Natal, Durban 2013.

Sounding rockets find applications in multiple fields of scientific research including meteorology, astronomy and microgravity. Indigenous sounding rocket technologies are absent on the African continent despite a potential market in the local aerospace industries. The UKZN Phoenix Sounding Rocket Programme was initiated to fill this void by developing inexpensive medium altitude sounding rocket modeling, design and manufacturing capacities. This dissertation describes the development of the Hybrid Rocket Performance Simulator (HYROPS) software tool and its application towards the structural design of the reusable, 10 km apogee capable Phoenix-1A hybrid sounding rocket, as part of the UKZN Phoenix programme. HYROPS is an integrated 6–Degree of Freedom (6-DOF) flight performance predictor for atmospheric and near-Earth spaceflight, geared towards single-staged and multi-staged hybrid sounding rockets. HYROPS is based on a generic kinematics and Newtonian dynamics core. Integrated with these are numerical methods for solving differential equations, Monte Carlo uncertainty modeling, genetic-algorithm driven design optimization, analytical vehicle structural modeling, a spherical, rotating geodetic model and a standard atmospheric model, forming a software framework for sounding rocket optimization and flight performance prediction. This framework was implemented within a graphical user interface, aiming for rapid input of model parameters, intuitive results visualization and efficient data handling. The HYROPS software was validated using flight data from various existing sounding rocket configurations and found satisfactory over a range of input conditions. An iterative process was employed in the aerostructural design of the 1 kg payload capable Phoenix-1A vehicle and CFD and FEA numerical techniques were used to verify its aerodynamic and thermo-structural performance. The design and integration of the Phoenix-1A‟s hybrid power-plant and onboard electromechanical systems for recovery parachute deployment and motor oxidizer flow control are also discussed. It was noted that use of HYROPS in the design loop led to improved materials selection and vehicle structural design processes. It was also found that a combination of suitable mathematical techniques, design know-how, human-interaction and numerical computational power are effective in overcoming the many coupled technical challenges present in the engineering of hybrid sounding rockets.
Thesis (M.Sc.)-University of KwaZulu-Natal, Durban, 2012.

碩士
國立成功大學
航空太空工程學系
104
For the past decades, aerospace exploration has gotten lots of attentions, and the authorities of Taiwan also spent great efforts on autonomous developments of aerospace techniques. For the mission requirements in the future, a low cost, reliable and safety “green” propellant is the first priority. On the use of upper stage rockets and kick motors, liquid rockets is always a better options because of its high specific impulse, controllable thrust and multi-impulse, but the injection and system design is too complicated and expensive. Without the supports of advanced technologies, it nearly impossible to achieve the goal in the next few years. Hybrid rockets definitely a good choice to replace liquid rockets with its advantages, like controllable thrust, simpler constructions and high safety. For the last decade, the NCKU hybrid rocket team has succeed launching few N2O hybrid rockets, with thrust level from 100 kgf to 3000 kgf, and our team has also developed a composite silver catalyst used on the H2O2 monopropellant thruster, which has an excellent performance. Therefore, this thesis is going to demonstrate a pre-decomposition H2O2 hybrid rocket using our catalyst’s techniques. Key words : hybrid rocket, HTP, silver catalyst INTRODUCTION For recent years, rocket-grade hydrogen peroxide propulsion systems got a renewed interest due to its low toxicity, low cost and minor impact to the environment. For the hydrogen peroxide concentration over 92% has been called High Test Peroxide(HTP), and the decomposition temperature of 100 wt% HTP can reach 1267K. Therefore, lots of researches was developed on HTP mono-propellant, and silver, manganese based catalyst were commonly used. Hybrid rocket technology is known for more than 50 years, its separation storage of fuels and oxidizers makes it safer than other propulsion systems. Nowadays, the need for green propellants, safe storability and the use of upper stage rockets made hybrid rockets more attractive. LOX, HTP and N2O are widely used as oxidizers due to their low toxicity and low pollutant characteristics. The initiate of hybrid rocket required a heat source to gasify fuels until reaching a combustible condition, while a ignition device is needed, this makes the construction heavier and more complicated. A pre-decomposition HTP hybrid rocket which can auto-ignite the fuel is developed, and several collages and research units have already gotten tremendous results. To shorten the period of development, hybrid rockets are definitely a good replacement, our team has lots of experiences on N2O hybrid rockets and also developed a composite silver catalyst using on the HTP mono-propellant, so this thesis is going to demonstrate an auto-ignition hybrid rocket using pre-decomposition HTP. ANALYST AND DESIGN The motor design is based on the 2000N thrust requirement of upper stage rockets or kick motors, with 1/8 reduced scale, this thesis is going to verify the preliminary test of a 250 N thrust engine. The CEA (Chemical Equilibrium with Application) code provided a starting point of this investigation. Assumed the combustion chamber pressure of 380 psi and using 90 wt% H2O2 as our oxidizer and 50P (50 wt% of HTPB+50 wt% of paraffin) fuels. From this data we predicted Isp and optimum O/F ratios as showed in figure 1. The O/F ratio of 7 was selected with the best predicted Isp of 240.6s, compared to the stoichiometric O/F ratio of 7.7, this selected ratio will lead the combustion process to go through a fuel rich combustion, which can prevent the nozzle from corrosion, with the figure we can calculate the other parameters. m ̇_total=T/(Isp×g) m ̇_fuel=m ̇_total/(1+O/F) From the equations above, we can get the mass flow rate of oxidizers and fuels. Catalyst Bed And Liquid HTP Injector Design The catalyst bed design is based on the 1N mono-propellant techniques our team has developed, to decompose the H2O2 of 92.75g/s, we used the data captured in the previous experiments, and going through a scale enlargement. The composite catalyst is composed of silver flakes and γ-Al2O3 pellets due to its stronger hardness. Before entering the catalyst bed, the H2O2 was designed to go through an injector to spread liquid H2O2 uniformly into the catalyst. m ̇=C_d A√2ρ∆P The equation above was used for injector design, after assuming the pressure drop of 15 psi and the physical properties, we can estimate the total area required. Therefore, the injector was designed with 16 bores, each bore has a diameter of 1 mm. Gaseous Injector Design Injector plays an important role in hybrid rocket mixing mechanism, decomposed H2O2 will be led into combustion chamber, mixing with gaseous fuels which gasified by high temperature gaseous H2O2, the whole process is thermal-related, to reduce the energy loss is our first consideration. The mass flow rate can be expressed as m ̇=ρ_2 U_2 A Where U_2、ρ_2 can be represented as U_2=√(2γ/((γ-1))×RT_1×[1-(〖p_2/p_1 )〗^((γ-1)/γ)]) ρ_2=ρ_1 〖[p_2/p_1 ]〗^(1/γ) After decomposing, the exit gas can be regarded as a mixture of O2 and gaseous H2O, assume ideal gas, and set p2/p1 as 0.9, with previous assumption of chamber pressure 380 psi, the pressure ahead of injector can be calculated. We can also calculate the value of density using the ideal gas equation of state, with the assumption of ideal decomposition temperature to be 1067 K and γ=1.26, the ideal exit velocity of 289.106 m/s was estimated. From the calculation, the injector was designed with 8 rectangle grooves each has a length of 5 mm and width of 2 mm. Fuel Grain NCKU rocket team has made a great effort on N2O hybrid rocket, and developed the 50P fuel, which composed of 50% paraffin and 50% HTPB, though 50P fuels was chosen as our solid fuels, with the data acquired from previous experiments, a regression rate of 1.5 mm/s was assumed. m_fuel=ρ×A×L With the specific weight of 0.9, we can estimate the dimension of the fuel, a hollow cylinder fuel grain was made, with the inner diameter of 20 mm, outside diameter of 53 mm and the length of 180 mm. Nozzle From CEA calculations, the total mass flow rate of 106 g/s and the reaction temperature 2700 K during combustion process was known, with γ=1.14 m ̇=p_c A_t γ√(〖[2/(γ+1)]〗^((γ+1)/(γ-1))/√(γRT_c )) With the equation above, a nozzle throat area of 64.615 mm2 graphite IG-11 was machined as our nozzle. RESULTS AND DISCUSSION Characteristics of Composite Silver Catalyst To verified the feasibility of auto-ignition, the attempt of using pre-decomposition to ignite the fuel was tried. The catalyst was designed and went through a scale enlargement from the 1N monopropellant we developed. Silver flakes and zeolite supports was used, but from the preliminary tests, zeolite was replaced byγ-Al2O3 due to its lack of hardness, and a terrific outcome was obtained. During decomposition process, the temperature over 800K was reached, and the catalyst chamber pressure was maintained steady through whole process, and the structure of the catalyst bed still remain stable after tests. Combustion Characteristics of 50P Fuels The motor was successfully auto-ignited in the tests, and from one of the tests listed below, the average thrust of 205 N was measured, the H2O2 flow rate of 92.05g and average chamber pressure of 352.18 psi was also measured in that test, from the figure below, a stable curvature was obtained. And from the definition of specific impulse, a value of 192.12s was reached in case 4. Isp=F ̅/(m ̇_f+m ̇_o ) Swirl Effects in Combustion Chamber From the earlier studies, applied swirl injection to the hybrid rocket would increase the performance of combustion efficiency, and the intensity of swirling was defined as the swirl number S=2/3((1-〖(R_h/R)〗^3)/(1-〖(R_h/R)〗^2 ))tanα Thus a swirl injector was applied, with different swirl number, the table listed below shows the data measured in the tests, with stronger swirling injection, the combustion efficiency was contrarily decreased, these results differed from the studies, and from the flame observation, the causes was concluded. Due to insufficient gasified fuels, the oxidizer/fuel mixing was poor, the un-gasified fuels was carried to the downstream of combustion chamber. CONCLUSION This thesis was tend to demonstrate the auto-ignition of hybrid rocket using decomposition H2O2, a thrust level of 250 N was set, and the swirling injection was applied as well, few results we got was listed below: Development of high capacity catalyst bed: γ-Al2O3 was used as catalyst support, and zeolite was replaced due to its weaker hardness. During the tests, the catalyst was able to sustain the high H2O2 flow rate and the high temperature. Establishment of firing system and procedure: This thesis was dedicated in establishing a procedure and method on auto-firing a H2O2 hybrid rocket, and successfully conform the design of H2O2 catalyst bed and hybrid rocket motor. Succeed auto-firing a hybrid rocket motor: During tests, the time delay of auto-firing are less than 0.08 second, and succeed producing thrust. The key on promoting the mixing of oxidizer and fuels: In the tests, all combustion conditions were in fuel rich, but from the observation of internal ballistics, a low combustion efficiency condition was observed. With the use of swirling injection, from the previous studies, it should improve the combustion efficiency. In our tests, the regression rate did improved, but the internal ballistics and the flame still shows a low combustion efficiency, this may due to low viscosity of paraffin based fuels, and cause an entrainment effect.

This chapter briefly introduces hybrid rocket propulsion for general audience. Advantageous of hybrid rockets over solids and liquids are presented. This chapter also explains how to design a test setup for hybrid motor firings. Hybrid propulsion provides sustainable, safe and low cost systems for space missions. Therefore, this chapter proposes hybrid propulsion system for Mars Ascent Vehicles. Paraffin wax is the fuel of the rocket. Propulsion system uses CO2/N2O mixture as the oxidizer. The goal is to understand the ignition capability of the CO2 as an in-situ oxidizer on Mars. CO2 is known as major combustion product in the nature. However, it can only burn with metallic powders. Thus, metallic additives are added in the fuel grain. Results show that CO2 increase slows down the chemical kinetics thus reduces the adiabatic flame temperature. Maximum flammability limit is achieved at 75% CO2 by mass in the oxidizer mixture. Flame temperature is 1700 K at 75% CO2. Ignition quenches below the 1700 K.

Abstract This research established methods of physical testing and analysis to characterize damage to a lander caused by rocket-propelled regolith during powered operations on or near the surface of the Moon, Mars, and other worlds. An emphasis was placed on using commonly available test materials to help lower cost. Lift-off using different thrusts were conducted in atmosphere from simulated planetary regolith. Testing was conducted with a scale model of a recent commercial lander with functioning scale rocket analogues using compressed gas as a propellant. This enabled testing to be safer and more cost-effective compared to operating rockets that burned fuel. Methods of quantifying and interpreting potential impact and abrasion damage and dust covering were established using crushable foam plates, adhesive coverings, and selective color processing of test result images. This testing can be an important and valuable step in a wider testing program to identify key design revisions before incurring the expense and risk to equipment from testing rocket and regolith interactions in a vacuum facility. This testing requires an awareness of the effects of scale and the limits of testing in atmosphere compared to the lower pressure or vacuum at the intended landing site.

The student team Hybrid Engine Development (HyEnD) of the University of Stuttgart is taking part with the Institute of Space Systems (IRS) in the DLR educational program STERN (Studentische Experimentalraketen). This program supports students at German universities to design, build, and launch an experimental rocket within a 3-year project time frame. HyEnD is developing a hybrid rocket called HEROS (Hybrid Experimental Rocket Stuttgart) with a design thrust of 10 kN, a total impulse of over 100 kN·s, and an expected liftoff weight up to 175 kg. HEROS is planned to be launched in October 2015 from Esrange in Sweden to an expected flight altitude of 40 to 50 km. The current altitude record for amateur rockets in Europe is at approximately 21 km. The propulsion system of HEROS is called HyRES (Hybrid Rocket Engine Stuttgart) and uses a paraffin-based solid fuel and nitrous oxide (N2O) as a liquid oxidizer. The development and the test campaign of HyRES is described in detail. The main goals of the test campaign are to achieve a combustion efficiency higher than 90% and provide stable operation with low combustion chamber pressure fluctuations. The successful design and testing of the HyRES engine was enabled by the evaluation and characterization of a small-scale demonstrator engine. The 500-newton hybrid rocket engine, called MIRAS (MIcro RAkete Stuttgart), has also been developed in the course of the STERN project as a technology demonstrator. During this test campaign, a ballistic characterization of paraffin-based hybrid rocket fuels with different additives in combination with N2O and a performance evaluation were carried out. A wide range of operating conditions, fuel compositions, injector geometries, and engine configurations were evaluated with this engine. Effects of different injector geometries and postcombustion chamber designs on the engine performance were analyzed. Additionally, the appearance of combustion instabilities under certain conditions, their effects, and possible mitigation techniques were also investigated. Concluding, the development and construction of an advanced, lightweight hybrid sounding rocket for the given requirements and budget within the DLR STERN program are described herein. The most important parts include a high thrust hybrid rocket engine, the development of a light weight oxidizer tank, pyrotechnical valves, carbon fiber rocket structure, recovery systems, and onboard electronics.