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Gas Power Cycles
Published in Kavati Venkateswarlu, Engineering Thermodynamics, 2020
Turbine and propeller engines use air from atmosphere, while rocket engines contain oxygen (oxidizer) within itself. Fuel and oxidizer are mixed and exploded in the combustion chamber, and the hot exhaust gases produced during combustion pass through the nozzle to accelerate the flow and produce the thrust required to propel the rocket. Turbine and propeller engines cannot operate in outer space since there is no atmosphere, but a rocket works in space. There are basically two types of rocket engines: liquid propellant and solid propellant. In the former case, the propellants (fuel and oxidizer) are stored separately as liquid and are sent to the combustor of the nozzle in which combustion takes place. In the latter case, the propellants are mixed together and packed into a solid cylinder. Burning takes place when propellants are exposed to a source of heat supplied by an igniter. The burning continues until all the propellants are exhausted. Due to the pumps and storage tanks, liquid propellant rockets tend to be heavier and complex compared to that of solid propellant that can be handled with ease.
Infrared Sources
Published in Monroe Schlessinger, Infrared Technology Fundamentals, 2019
Most rocket engines are designed to maximize the specific impulse (Isp) attained for the propellants used. This involves both the rocket thrust produced and the rate of flow of combustion gases through the nozzle throat and eventually out of the nozzle. As a result, in most rocket engines the products of combustion usually contain substances (such as H2 and CO) that are still combustible. When these reach the atmosphere, they combine with the available oxygen to create afterburning in the exhaust plume.
Project in Rocketry
Published in G. Boothroyd, C. Poli, Applied Engineering Mechanics, 2018
A rocket engine produces gas molecules which are ejected with a high velocity; this causes a reaction which propels the rocket forward. To predict the motion of the rocket, it is necessary to obtain an equation which relates the forces acting on the rocket to its instantaneous acceleration. Forces are generally not applied to objects at a point but are distributed in their effect over a finite area. Under these circumstances, it is more useful to refer to a pressure; this is a force divided by the area over which the force is exerted. One effect of the ejection of gas molecules from a rocket engine is to produce a pressure at the rocket engine nozzle which helps to propel the rocket forward. The effect of this pressure is reduced to some extent by the pressure exerted by the atmosphere (atmospheric pressure). Thus, the effective pressure at the nozzle is found by subtracting the atmospheric pressure from the pressure of the exhaust gases. The force Fp resulting from this pressure is then found by multiplying the effective pressure p by the cross-sectional area A of the rocket nozzle.
Three-dimensional printed metal-nested composite fuel grains with superior mechanical and combustion properties
Published in Virtual and Physical Prototyping, 2022
Xin Lin, Dandan Qu, Xuedong Chen, Zezhong Wang, Jiaxiao Luo, Dongdong Meng, Guoliang Liu, Kun Zhang, Fei Li, Xilong Yu
Hybrid rocket engines (HREs) combine the intrinsic advantages of liquid propellants and solid fuels, which renders it simple structure, high safety and reliability, adjustable thrust, and lower cost than conventional rocket engines (Whitmore, Sobbi, and Walker 2014; Wang et al. 2021a; Cai et al. 2013; Fang et al. 2021; Zilliac et al. 2020; Kahraman, Ozkol, and Karabeyoglu 2021). These advantages make HREs attractive with regard to a broad range of space applications, such as sounding rockets (Sella et al. 2020; Bouziane et al. 2019; Broughton et al. 2018; Marciniak et al. 2018), upper stage propulsion units (Jens, Cantwell, and Hubbard 2016; Casalino and Pastrone 2008) and commercial manned spacecrafts (Cai et al. 2013; Mazzetti, Merotto, and Pinarello 2016). However, HREs have several associated challenges, the most significant being the low regression rates of classical polymeric fuels. The issue of low regression rates has thus far seriously restricted the implementation of HRE technology in large-scale thrust applications (Kobald et al. 2017).
Computational investigation of cooling effectiveness for film cooled dual-bell exhaust nozzle for LO2/LH2 liquid rocket engines
Published in Energy Sources, Part A: Recovery, Utilization, and Environmental Effects, 2021
Martin Raju, Abhilash Suryan, David Šimurda
Specific impulse () is the thrust generated by a rocket engine per weight flow rate of fuel consumed. Specific impulse loss due to secondary coolant injection is due to the excess fuel consumption used to cool the nozzle walls. As the coolant is injected at the inflection, which is situated downstream from the inlet of DBN, it does not take part in the combustion process unlike in case of combustion chamber cooling (Batha et al. 1963). Specific impulse is calculated using the equation given below: