Begell House Inc.
International Journal of Energetic Materials and Chemical Propulsion
IJEMCP
2150-766X
16
2
2017
ADVANCED ENERGETIC MATERIALS AND TECHNIQUES FOR ROCKET PROPULSION: IN HONOR OF PROF. KENNETH KUAN-YUN KUO
v-xiii
10.1615/IntJEnergeticMaterialsChemProp.v16.i2.10
Luigi T.
De Luca
Space Propulsion Laboratory (SPLab), Department of Aerospace Science
and Technology, Politecnico di Milano, I-20156 Milan, Italy
Keiichi
Hori
Institute of Space and Astronautical Science (ISAS), Japan Aerospace Exploration
Agency (JAXA), Sagamihara, Kanagawa, 252-5210, Japan
John
Zevenbergen
TNO Defense Security and Safety
EFFECT OF AZODICARBONAMIDE PARTICLES ON THE REGRESSION RATE OF HYDROXYL-TERMINATED POLYBUTADIENE (HTPB)-BASED FUELS FOR HYBRID ROCKET PROPULSION
103-114
10.1615/IntJEnergeticMaterialsChemProp.2018022374
Suhang
Chen
School of Chemical Engineering, Nanjing University of Science and Technology,
Nanjing, Jiangsu, 210094, China
Yue
Tang
School of Chemical Engineering, Nanjing University of Science and Technology,
Nanjing, Jiangsu, 210094, China
Wei
Zhang
School of Chemical Engineering, Nanjing University of Science and Technology
200 Xiaolingwei, Xuanwu District, Nanjing 210094, China
Ruiqi
Shen
Nanjing University of Science and Technology
Luigi
DeLuca
Politecnico di Milano
Yinghua
Ye
Institute of Space Propulsion, School of Chemical Engineering, Nanjing
University of Science and Technology, Nanjing, 210094, China
hybrid propulsion
regression rate
ADCA/HTPB fuels
porous layer
The details of the burning surface layer can have a great influence on the combustion performance of solid fuels for hybrid propulsion. In composite solid fuels, a substance with a low decomposition temperature or a fast burning rate leaves holes in the initial matrix after decomposition or burning, thus increasing the regressing surface area. Azodicarbonamide (ADCA)/hydroxyl-terminated
polybutadiene (HTPB) composite fuels can achieve this kind of mechanism: ADCA decomposes at the temperature of 150–306°C while HTPB decomposes at 417–591°C, thus forming a certain thickness of porous layer after the decomposition of ADCA particles when the fuel burns. However, the gaseous products increase the blocking effect, hindering heat transfer from the flame zone to the fuel surface and suppressing combustion. In this study, we investigate the effects of ADCA on the combustion characteristics of HTPB-based composite solid fuels. The first exothermic peak of thermal gravimetry and differential scanning calorimetry (TG-DSC) tests for ADCA/HTPB fuels decreases, while the second exothermic peak increases in intensity and simultaneously shifts to lower temperatures with increasing ADCA contents. The optimal addition to the HTPB matrix of 3% in mass of ADCA particles enhanced the pre-expansion during combustion and showed an increased regression rate by 41.52%, while 1% and 5% in mass of ADCA increased by 2.99% and –0.40%, respectively, at GOX = 340 kg/m2s, but revealed a decrease to –15.61%, –11.81%, and –25.74%, correspondingly at GOX = 150 kg/m2s. This suggests that appropriate ADCA acts as an effective burning rate modifier by favoring the formation of a certain thickness of porous layer.
IGNITION AND COMBUSTION CHARACTERISTICS OF A MICRO-ELECTROMECHANICAL SYSTEM (MEMS) PYROTECHNIC THRUSTER FOR MICRO PROPULSION APPLICATIONS
115-123
10.1615/IntJEnergeticMaterialsChemProp.2018024455
Harshit
Shukla
Department of Space Engineering and Rocketry, Birla Institute of Technology,
Mesra, Ranchi, 835215, India
Gusain Rajesh Singh
Nandan
Vikram Sarabhai Space Centre, ISRO, Thiruvananthapuram, 695022, India
Priyanka
Shukla
Vikram Sarabhai Space Centre, ISRO, Thiruvananthapuram, 695022, India
Vinod
Kumar
Vikram Sarabhai Space Centre, ISRO, Thiruvananthapuram, 695022, India
Mohan
Varma
Department of Space Engineering and Rocketry, Birla Institute of Technology, Mesra, Ranchi, 835215, India
energetic materials
solid propellant thruster
micro-spacecraft propulsion
micro-igniters
heater configuration
A functional micro-electromechanical system (MEMS)-based micro-thruster using boron–potassium nitrate pyrotechnic composition with an electron beam deposited platinum heater has been characterized for reproducible ignition and sustained combustion. Configuration of the device,
assembly process, and the energetic material were selected based on a detailed study involving a series of ignition tests. Successful ignition was achieved for the chosen configuration and the energetic material. The development demonstrated potential of such devices along with scalability for futuristic applications aimed toward downsizing of micro-spacecraft propulsion systems.
EFFECTS OF DIFFERENT DECELERATION AGENTS ON THE PROPERTIES OF HYDROXYL TERMINATED POLYETHER (HTPE)-BASED COMPOSITE SOLID PROPELLANTS
125-138
10.1615/IntJEnergeticMaterialsChemProp.2018025006
Wei Qiang
Pang
Xi'an Modern Chemistry Research Institute, Xi'an, Shaanxi, China
Jun Qiang
Li
Xi'an Modern Chemistry Research Institute, Xi'an, Shaanxi, 710065, People's
Republic of China
Luigi
DeLuca
Politecnico di Milano
Ke
Wang
Xi'an Modern Chemistry Research Institute, Xi'an, Shaanxi, 710065, People's
Republic of China
XiaoLong
Fu
Xi'an Modern Chemistry Research Institute, Xi'an, Shaanxi, 710065, People's
Republic of China
Xue Zhong
Fan
Xi'an Modern Chemistry Research Institute, Xi'an, Shaanxi, 710065, People's
Republic of China
Huan
Li
Xi'an Modern Chemistry Research Institute, Xi'an, Shaanxi, 710065, People's
Republic of China
material chemistry
HTPE
composite solid propellant
deceleration agents
hazardous properties
combustion characteristics
Several industrial- and research types of hydroxyl-terminated polyether (HTPE) composite solid
propellants containing different deceleration agents, featuring the same nominal composition, were
prepared and experimentally analyzed. The energetic properties and the hazardous properties of propellants prepared were analyzed. The effects of different deceleration agents on the strand burning rate and the associated combustion flame structure of propellants were investigated and the properties mentioned above compared to that of a conventional composite propellant without deceleration agents, and the thermal decomposition of propellants were performed. It was shown that the application of different burning rate suppressant additives in the HTPE-based composite solid propellants is feasible, which can be casted in vacuum and cured safely. The HTPE solid propellants containing various deceleration agents are insensitive to impact and friction, and the heat of explosion of reference composite solid propellant ([CP] sample) is higher than those of the propellants with different deceleration agents. Moreover, the addition of deceleration agents to the propellant formulations can decrease the burning rates compared to the reference propellant without any deceleration agent. Among the composite solid propellant formulations, the pressure exponent of the propellant containing di-butyl phthalate ([DB-CP] sample) is the lowest one (0.212).
IGNITION AND COMBUSTION OF HYDROXYL-TERMINATED POLYBUTADIENE (HTPB)-BASED SOLID FUELS LOADED WITH INNOVATIVE MICROMETER-SIZED METALS
139-150
10.1615/IntJEnergeticMaterialsChemProp.2018022772
Zhao
Qin
Science and Technology on Combustion and Explosion Laboratory, Xi'an Modern Chemistry Research Institute, No. 168 Zhangbadonglu, Yanta District, Xi'an, 710065, China; Nanjing University of Science and Technology, Chemical Engineering School, No. 200 Xiaolingwei Street, Xuanwu District, Nanjing, 210094, China
Christian
Paravan
Politecnico di Milano
Giovanni
Colombo
Politecnico di Milano, Department of Aerospace Science and Technology, Via La
Masa 34, Milan, 20133, Italy
Feng-qi
Zhao
Science and Technology on Combustion and Explosion Laboratory, Xi'an
Modern Chemistry Research Institute, No. 168 Zhangbadonglu, Yanta District,
Xi'an, 710065, China
Ruiqi
Shen
Nanjing University of Science and Technology
Jian-hua
Yi
Science and Technology on Combustion and Explosion Laboratory, Xi'an
Modern Chemistry Research Institute, No. 168 Zhangbadonglu, Yanta District,
Xi'an, 710065, China
Luigi
DeLuca
Politecnico di Milano
ignition delay
regression rate
combustion
pressure
micrometer-sized metal
This paper discusses experimental investigations on the effects of innovative micrometer-sized metal additives on the ignition and burning of solid fuel formulations based on Hydroxyl-Terminated PolyButadiene (HTPB). Gaseous oxygen was selected as the oxidizer for ignition delay and regression rate tests. A relative grading of solid fuel performance was carried out taking unloaded HTPB as the baseline. Three different micrometer-sized metallic additives were investigated: conventional magnesium (Mg), amorphous aluminum (am_Al,) and magnesium–boron composite (MgB). Ignition delay is highly depending on pressure, while the linear regression rate was not appreciably affected as the pressure increased from 1.0 to 1.9 MPa. All of the micrometer-sized additives have a positive effect on enhancing both linear regression rate and mass burning rate, while amorphous aluminum (am_Al) demonstrated a larger effect than Mg and composite MgB powders.
CYCLIC QUENCHING AND IGNITION OF LOW-STRETCH DIFFUSION FLAMES
151-163
10.1615/IntJEnergeticMaterialsChemProp.2018022944
Gary P.
Garbinski
Department of Mechanical and Aerospace Engineering, Case Western Reserve
University, Cleveland, Ohio 44106, USA
James S.
T'ien
Department of Mechanical and Aerospace Engineering, Case Western Reserve University, Cleveland, OH 44106, USA
diffusion flame
quenching
ignition
low stretch
cyclic phenomena
oscillation
A transient numerical simulation on the response of counter-flow gaseous diffusion flames was carried out over the entire range of stretch rates with the fuel and oxygen ambient temperatures as varying parameters. The model assumes an overall chemical reaction but with flame radiation loss. The shapes of steady flame response curve obtained include "isola" at low ambient temperatures
and "ghost" at high ambient temperatures. A cyclic flame ignition and quenching phenomenon was discovered in a low stretch region with moderate ambient temperatures. This paper presents the computed response curves and, in particular, the time-resolved cyclic event. In addition, a map using ambient temperature and stretch rate as coordinates is given to show the various regions of
the diffusion flame response.
HIGH-TEMPERATURE COATING: HYBRID ROCKET MOTOR THERMAL PROTECTION CASE HISTORY
165-174
10.1615/IntJEnergeticMaterialsChemProp.2018023565
Marcela Galizia
Domingues
ITA – Instituto Tecnológico de Aeronáutica, Departamento de Química, Praça
Marechal Eduardo Gomes, 50 - Vila das Acácias, São José dos Campos - SP,
12228-900, Brazil
José Atílio Fritz Fidel
Rocco
ITA – Instituto Tecnológico de Aeronáutica, Departamento de Química, Praça
Marechal Eduardo Gomes, 50 - Vila das Acácias, São José dos Campos - SP, 12228-900, Brazil
high-temperature coating
thermal protection
hybrid rocket motor
water glass
silicon carbide
Most of the coatings available on the market do not meet the protection requirements of high temperatures and oxidation of metallic surfaces exposed to chemically aggressive environments at high temperatures. Thus, the development of an inorganic base coating (potassium silicate) was proposed
in the form of an aqueous solution for application in hybrid rocket motor metallic components. This coating distinguishes itself from others by supporting extreme operating conditions of operation without degrading or losing its original characteristics, thus forming a glassy film which anchors in/on the surface of the substrate in which it is applied, i.e., becoming a superficial layer. The investigation of the applicability of a high-temperature coating was studied in adverse conditions, as in
the case of a hybrid rocket motor; different components of the engine have received the coating application.
Once coated, the components were assembled as a hybrid engine and subjected to firing tests. The results were very promising, since the coating could reduce the erosion in the throat of the nozzle by 50%, improving the hybrid rocket motor operation time at temperatures around 2.000 Celsius.
COUPLING OF TRANSIENT THERMAL AND MECHANICAL STRESSES COMPUTATIONS IN GRAPHITE NOZZLE MATERIALS
175-195
10.1615/IntJEnergeticMaterialsChemProp.2018024875
Ragini
Acharya
University of Tennessee Space Institute
Brian
Evans
Pennsylvania State University, University Park, Pennsylvania 16802, USA
Jonathan
Pitt
Pennsylvania State University, University Park, Pennsylvania 16802, USA
Francesco
Costanzo
Pennsylvania State University, University Park, Pennsylvania 16802, USA
Kenneth K.
Kuo
Department of Mechanical and Nuclear Engineering, The Pennsylvania State University, University Park, PA 16802, USA
ignition transient
graphite
nozzle
thermal stresses
solid rocket nozzle
numerical simulation
In this work, a numerical simulation of the structural response of the graphite nozzle materials to the flow during the ignition transient of a solid rocket motor is considered. The measured pressure–time trace within the combustion chamber was used as an input parameter in the graphite-nozzle erosion minimization (GNEM) code to calculate gas-phase pressure, temperature, velocity, etc. in
the graphite nozzle. The calculated pressure and convective heat flux from GNEM were applied as loading conditions in an associated thermo-structural model to obtain response of graphite nozzle materials to the transient pressure and thermal loading. The combined aero–thermo–structural response of the graphite rocket nozzle showed that the thermal stresses were significantly higher than the mechanical stresses. The radial displacements of inner nozzle surface showed that the nozzle diameter increases in the beginning due to thermo-mechanical stresses. The axial displacements of several points on the inner nozzle surface showed that the surface tends to displace toward the entrance plane.