This work is focused on the numerical simulation and experimental characterization of a vortex flow pancake (VFP) hybrid rocket engine (HRE). In this rocket configuration, a tangential injector is sandwiched by two fuel disks. One of the two fuel disks has a port channel leading to the exhaust nozzle. During the combustion the propellant mixture mixing is enhanced by the vortex flow realized between the two disks thanks to the oxidizer tangential injection. This HRE configuration offers high compactness and an easy-fit to different platforms making the system attractive for implementation on different satellites for active de-orbiting missions. The first part of the work was devoted to an evaluation of the VFP HRE as a candidate for de-orbiting applications. In particular, this was studied analysing by Montecarlo Method the effect of uncertainties on spacecraft re-entry. Different systems were compared in the analysis: a solid rocket motor (SRM), a liquid rocket engine (LRE) and the VFP. In spite of the simplifying assumptions made for the re-entry problem, the LRE and the HRE systems featured attractive performance in the minimization of the spacecraft re-entry footprint, thanks to thrust throttling and burn duration control. Between these two systems, the VFP was selected in light of its intrinsic safety, simplicity and low-cost. Evaluation of three propellant sets was carried out, resulting in the choice of the nitrous oxide - hydroxyl-terminated polybutadiene (HTPB) as optimal one, granting further flexibility and simplicity in the system implementation. Therefore, in the second part of the work, nitrous oxide combustion behaviour with HTPB was investigated. Regression rate was measured for nitrous oxide mass flow rates of 5.5 g/s and 8.25 g/s. Throttling of the engine was performed, changing the oxidizer mass flow rate from 8.25 to 5.5 g/s. Engine cumulative firing time achieved was roughly four minutes, the pressure range investigated extended from 0.63 to 1.06 MPa. Wire cut sensors were implemented in the engine to retrieve the local, instantaneous regression rate. Given these considerations, VFP HRE shows attractive features for application in controlled de-orbiting missions.
Il presente lavoro è incentrato sulla simulazione numerica e la caratterizzazione sperimentale di un motore ibrido per razzi (HRE) con iniezione a vortice (VFP). In questa configurazione, l’iniezione dell’ossidante avviene tangenzialmente tra due dischi di combustibile. Durante la combustione, il miscelamento di ossidante e combustibile è favorito dal campo di moto vorticoso generato dall’iniezione tangenziale. Questa tipologia di HRE presenta un’ottima compattezza e facilità di integrazione per varie tipologie di piattaforme, rendendolo una scelta attraente per l’implementazione su vari satelliti per missioni di de-orbiting attivo. Nella prima parte di questo lavoro è stata implementata una simulazione numerica di una missione di de-orbiting in cui il motore utilizzato è il VFP. In particolare, è stato analizzato l’effetto delle incertezze sulla precisione del rientro atmosferico utilizzando il metodo Montecarlo. In questa analisi sono stati comparati un motore a propellente solido (SRM), uno a liquido (LRE) ed il VFP. Benchè siano state utilizzate delle ipotesi semplificative nella formulazione del problema, i motori VFP e LRE mostrano prestazioni interessanti per quanto riguarda la minimizzazione del footprint di rientro grazie alla possibilità di controllare la spinta, in termini sia di durata che di modulo. Tra questi due ultimi sistemi, il VFP è stato scelto come alternativa migliore grazie alla sua intrinseca sicurezza, semplicità architetturale e basso costo. Inoltre, sono state valutate tre possibili coppie ossidante - combustibile. La coppia protossido di azoto - polibutadiene a terminazione idrossilica (HTPB) è risultata la migliore delle tre, essendo capace di garantire ulteriore flessibilità e semplicità nel sistema. Di conseguenza, nella parte sperimentale è stata studiata la combustione di protossido di azoto con HTPB. Per tale coppia ossidante-combustibile è stato misurato il rateo di regressione della superficie di combustibile utilizzando portate massiche di protossido di azoto pari a 5.5 g/s e 8.25 g/s. In una prova è stata effettuata la regolazione della portata massica di ossidante, passando da 8.25 a 5.5 g/s. Il tempo cumulativo di combustione raggiunto è stato di circa quattro minuti, con pressioni variabili da 0.63 a 1.06 MPa. Sensori a rottura di filo sono stati montati sul motore in modo da ottenere il rateo di regressione istantaneo e locale. In conclusione, il VFP mostra caratteristiche interessanti per la sua implementazione in missioni di de-orbiting attivo.
A non-conventional hybrid rocket motor for de-orbiting applications
MARINO, FRANCESCO
2016/2017
Abstract
This work is focused on the numerical simulation and experimental characterization of a vortex flow pancake (VFP) hybrid rocket engine (HRE). In this rocket configuration, a tangential injector is sandwiched by two fuel disks. One of the two fuel disks has a port channel leading to the exhaust nozzle. During the combustion the propellant mixture mixing is enhanced by the vortex flow realized between the two disks thanks to the oxidizer tangential injection. This HRE configuration offers high compactness and an easy-fit to different platforms making the system attractive for implementation on different satellites for active de-orbiting missions. The first part of the work was devoted to an evaluation of the VFP HRE as a candidate for de-orbiting applications. In particular, this was studied analysing by Montecarlo Method the effect of uncertainties on spacecraft re-entry. Different systems were compared in the analysis: a solid rocket motor (SRM), a liquid rocket engine (LRE) and the VFP. In spite of the simplifying assumptions made for the re-entry problem, the LRE and the HRE systems featured attractive performance in the minimization of the spacecraft re-entry footprint, thanks to thrust throttling and burn duration control. Between these two systems, the VFP was selected in light of its intrinsic safety, simplicity and low-cost. Evaluation of three propellant sets was carried out, resulting in the choice of the nitrous oxide - hydroxyl-terminated polybutadiene (HTPB) as optimal one, granting further flexibility and simplicity in the system implementation. Therefore, in the second part of the work, nitrous oxide combustion behaviour with HTPB was investigated. Regression rate was measured for nitrous oxide mass flow rates of 5.5 g/s and 8.25 g/s. Throttling of the engine was performed, changing the oxidizer mass flow rate from 8.25 to 5.5 g/s. Engine cumulative firing time achieved was roughly four minutes, the pressure range investigated extended from 0.63 to 1.06 MPa. Wire cut sensors were implemented in the engine to retrieve the local, instantaneous regression rate. Given these considerations, VFP HRE shows attractive features for application in controlled de-orbiting missions.| File | Dimensione | Formato | |
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https://hdl.handle.net/10589/139533