scholarly journals Hybrid Rocket Engine Design Optimization at Politecnico di Torino: A Review

Aerospace ◽  
2021 ◽  
Vol 8 (8) ◽  
pp. 226
Author(s):  
Lorenzo Casalino ◽  
Filippo Masseni ◽  
Dario Pastrone
Keyword(s):  
Design Optimization ◽  
Rocket Engine ◽  
Propulsion System ◽  
Future Perspective ◽  
Rocket Engines ◽  
Hybrid Rocket ◽  
Feed System ◽  
Research Activities ◽  
Multi Stage ◽  
System Design Optimization

Optimization of Hybrid Rocket Engines at Politecnico di Torino began in the 1990s. A comprehensive review of the related research activities carried out in the last three decades is here presented. After a brief introduction that retraces driving motivations and the most significant steps of the research path, the more relevant aspects of analysis, modeling and achieved results are illustrated. First, criteria for the propulsion system preliminary design choices (namely the propellant combination, the feed system and the grain design) are summarized and the engine modeling is presented. Then, the authors describe the in-house tools that have been developed and used for coupled trajectory and propulsion system design optimization. Both deterministic and robust-based approaches are presented. The applications that the authors analyzed over the years, starting from simpler hybrid powered sounding rocket to more complex multi-stage launchers, are then presented. Finally, authors’ conclusive remarks on the work done and their future perspective in the context of the optimization of hybrid rocket propulsion systems are reported.

Conceptual Regenerative Nozzle Cooling Design for a Hydroxyl-Terminated Polybutadiene and Oxygen Hybrid Rocket Engine

2017 ◽  
Author(s):  
Luis R. Robles ◽  
Johnny Ho ◽  
Bao Nguyen ◽  
Geoffrey Wagner ◽  
Jeremy Surmi ◽  
...  
Keyword(s):  
Combustion Chamber ◽  
High Temperatures ◽  
Cooling System ◽  
Rocket Engine ◽  
Combustion Efficiency ◽  
Rocket Engines ◽  
Hybrid Rocket ◽  
Gaseous Oxygen ◽  
Rocket Nozzle ◽  
Regenerative Cycle

Regenerative rocket nozzle cooling technology is well developed for liquid fueled rocket engines, but the technology has yet to be widely applied to hybrid rockets. Liquid engines use fuel as coolant, and while the oxidizers typically used in hybrids are not as efficient at conducting heat, the increased renewability of a rocket using regenerative cycle should still make the technology attractive. Due to the high temperatures that permeate throughout a rocket nozzle, most nozzles are predisposed to ablation, supporting the need to implement a nozzle cooling system. This paper presents a proof-of-concept regenerative cooling system for a hybrid engine which uses hydroxyl-terminated polybutadiene (HTPB) as its solid fuel and gaseous oxygen (O2) as its oxidizer, whereby a portion of gaseous oxygen is injected directly into the combustion chamber and another portion is routed up through grooves on the exterior of a copper-chromium nozzle and, afterwards, injected into the combustion chamber. Using O2 as a coolant will significantly lower the temperature of the nozzle which will prevent ablation due to the high temperatures produced by the exhaust. Additional advantages are an increase in combustion efficiency due to the heated O2 being used for combustion and an increased overall efficiency from the regenerative cycle. A computational model is presented, and several experiments are performed using computational fluid dynamics (CFD).

scholarly journals Burn Time Correction of Start-Up Transients for CAMUI Type Hybrid Rocket Engine

Aerospace ◽  
2021 ◽  
Vol 8 (12) ◽  
pp. 385
Author(s):  
Tor Viscor ◽  
Hikaru Isochi ◽  
Naoto Adachi ◽  
Harunori Nagata
Keyword(s):  
Rocket Engine ◽  
State Level ◽  
Impinging Jet ◽  
Rocket Engines ◽  
Transient Effects ◽  
Hybrid Rocket ◽  
Start Up ◽  
Time Correction ◽  
Hybrid Rocket Engine ◽  
Time Issue

Burn time errors caused by various start-up transient effects have a significant influence on the regression modelling of hybrid rockets. Their influence is especially pronounced in the simulation model of the Cascaded Multi Impinging Jet (CAMUI) hybrid rocket engine. This paper analyses these transient burn time errors and their effect on the regression simulations for short burn time engines. To address these errors, the equivalent burn time is introduced and is defined as the time the engine would burn if it were burning at its steady-state level throughout the burn time to achieve the measured total impulse. The accuracy of the regression simulation with and without the use of equivalent burn time is then finally compared. Equivalent burn time is shown to address the burn time issue successfully for port regression and, therefore, also for other types of cylindrical port hybrid rocket engines. For the CAMUI-specific impinging jet fore-end and back-end surfaces, though, the results are inconclusive.

scholarly journals ОСОБЛИВОСТІ КОНСТРУКТИВНИХ СХЕМ ДВИГУНІВ З ІМПУЛЬСНИМИ ДЕТОНАЦІЙНИМИ КАМЕРАМИ

2020 ◽  
pp. 4-10
Author(s):  
Олександр Євгенович Золотько ◽  
Олена Василівна Золотько ◽  
Олександра Валеріївна Сосновська ◽  
Олександр Сергійович Аксьонов ◽  
Ірина Сергіївна Савченко
Keyword(s):  
Electric Power ◽  
Propulsion System ◽  
High Rate ◽  
Rocket Engines ◽  
Pulsation Frequency ◽  
Feed System ◽  
Jet Engine ◽  
Pulse Detonation ◽  
Pulsed Detonation ◽  
Pulsed Detonation Engines

The pressure of the products of chemical reactions in the chamber of a rocket engine increases significantly if the rocket fuel components burn in the detonation mode. In this case, it can get to a simpler and more reliable expulsion propellant feed system instead of a turbopump feed system. The value of heat release power (MW / liter) of detonation engines is several orders of magnitude larger than that of aircraft and rocket engines operating in the Brighton cycle. The high rate of energy released in the detonation mode can significantly reduce the mass, the inertia, and overall dimensions of the propulsion system. Due to these features, detonation chambers are advisable to be used as part of ejector pulsed detonation engines, together with a turbine – in electric power generators of spacecraft, in a hybrid design – together with turbofan or turboprop engines, etc. In the article are considered various design schemes of pulse detonation engines (PDE): single-chamber and multi-chamber pulsed detonation engines; an ejector PDE system; a hybrid PDE and an integrated detonation-turbine unit with a detonation chamber in the form of a spiral and with a multi-chamber detonation device. The possibility of pulsation frequency increase is realized in the multi-chamber pulsed detonation engine, and the possibility of thrust size increase is realized in PDE with ejector. Replacing traditional chambers with detonation chambers in the construction of gas turbine jet engine will allow providing a decrease in propellant flow rate value from 8 % to 10 % on some estimations. In the hybrid detonation propulsion plant advantages inherent to the detonation cycle combine with positive features of a turbo-compressor jet engine. A combination of PDE and turbine allows creating the cogeneration propulsion system in that a turbine is used for the production of electric power, and detonation chamber – for the creation of thrust impulse. Practical realization of hybrid pulse detonation turbo-engine and the integrated detonation-turbine device is possible if two key complex problems will be solved. These problems are the detonation waves weakening on input in a turbine and the bearing and shaft necessary work resource increasing into a detonation pulsating stream

scholarly journals Comparison of cylindrical and non-cylindrical grain internal ballistic behavior of hybrid rocket engines and solid rocket motors

2021 ◽  
Author(s):  
Zhiliu Lu
Keyword(s):  
Specific Impulse ◽  
Propulsion System ◽  
Simulation Program ◽  
Rocket Engines ◽  
Hybrid Rocket ◽  
Ballistic Performance ◽  
Flow Process ◽  
Rocket Motors ◽  
Internal Ballistic ◽  
Propellant Rocket

Hybrid rocket engines (HREs) are a chemical propulsion system that nominally combine the advantages of liquid-propellant rocket engines (LREs) and solid-propellant rocket motors (SRMs). HREs in some cases can have a higher specific impulse and better controllability than SRMs, and lower cost and engineering complexity than LREs. For HREs and SRMs, both kinds of rocket engine employ a solid fuel grain, and the chosen grain configuration is a crucial point of their design. Different grain configurations have different internal ballistic behavior, which in turn can deliver different engine performance. A cylindrical grain design is a very common design for SRMs and HREs. A non-cylindrical-grain is a more complex grain configuration (than cylindrical) that has been used in many SRMs, and is also a choice for some HREs. However, while an HRE and an SRM can employ the same fuel grain configuration, the resulting internal ballistic behavior would not be expected to be the same. Pressure-dependent burning tends to dominate in SRMs, while axial flow-dependent burning tends to dominate in HREs. To help demonstrate in a more direct manner the influence of the differing combustion processes on the same fuel grain configuration used by an HRE and SRM, a number of internal ballistic simulations are undertaken for the present study. For the reference SRM cases looked at, an internal ballistic simulation program that has the capability of predicting head-end pressure and thrust as a function of time into a simulated firing is utilized for the present investigation; for the corresponding HRE cases, a simulation program is used to simulate the burning and flow process of these engines. For the present investigation, the two simulation programs are used to simulate the internal ballistic performance of various HREs and SRMs employing comparable cylindrical and non-cylindrical fuel grain configurations. The predicted performance results, in terms of pressure and thrust, are consistent with expectations that one would have for both propulsion system types.

scholarly journals Design, Production and Evaluation of 3D-Printed Mold Geometries for a Hybrid Rocket Engine

Aerospace ◽  
2021 ◽  
Vol 8 (8) ◽  
pp. 220
Author(s):  
Benedict Grefen ◽  
Johannes Becker ◽  
Stefan Linke ◽  
Enrico Stoll
Keyword(s):  
Solid Fuel ◽  
Degrees Of Freedom ◽  
Rocket Engine ◽  
Mold Material ◽  
Rocket Engines ◽  
Test Bed ◽  
Hybrid Rocket ◽  
Design Production ◽  
3D Printed ◽  
Hybrid Rocket Engine

The feasibility of 3D-printed molds for complex solid fuel block geometries of hybrid rocket engines is investigated. Additively produced molds offer more degrees of freedom in designing an optimized but easy to manufacture mold. The solid fuel used for this demonstration was hydroxyl-terminated polybutadiene (HTPB). Polyvinyl alcohol (PVA) was chosen as the mold material due to its good dissolving characteristics. It is shown that conventional and complex geometries can be produced reliably with the presented methods. In addition to the manufacturing process, this article presents several engine tests with different fuel grain geometries, including a short overview of the test bed, the engine and first tests.

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