An investigation of injectors for use with high vapor pressure propellants with applications to hybrid rockets
Abstract/Contents
- Abstract
- Interest in nitrous oxide based hybrid rockets is at an all time high. Nitrous oxide (N2O) is a unique oxidizer because it exhibits a high vapor pressure at room temperature (≈ 730 psia or 5.03 MPa). Due to this high vapor pressure, liquid nitrous oxide can be expelled from a tank without the use of complicated pumps or pressurization systems required by most traditional liquid rocket systems. This results in weight savings and design simplicity. Additional benefits of nitrous oxide include storability, ease of handling, and relative safety compared to traditional liquid oxidizers. The design and modeling of injectors for use with high vapor pressure propellants such as nitrous oxide is made complicated due to the possibility of two-phase flow. The operating pressures within rocket propellant feed systems can often drop below the vapor pressure for these unique propellants, especially within the injector. Injectors operating under these conditions are likely to exhibit cavitation, resulting in significant vapor formation and limitation of mass flow rate. A variety of two-phase flow models which attempt to account for this phenomenon are presented and discussed, however none have proven reliable enough to replace traditional experimental injector flow studies. For this reason, a small scale experimental injector cold flow rig was designed, and a test campaign was carried out in an effort to characterize the mass flow rate performance of nitrous oxide rocket injectors over a broad range of operating conditions. Some representative results from this campaign are presented. The Peregrine Sounding Rocket is a hybrid rocket that runs on paraffin wax and nitrous oxide. Its development is a joint effort between NASA Ames Research Center, Stanford University, and Space Propulsion Group, Inc. For years, progress of the Peregrine program has been hampered by combustion instability problems. Based upon results from the aforementioned small scale injector experiments, a powerful, yet simple solution to the so-called feed system coupled combustion instability was discovered, the details of which are presented. This work also led to the invention of a new class of rocket propellant injectors designed specifically to decrease the likelihood of this type of combustion instability. An in-depth discussion of the proposed design and operation of this novel injection scheme is included, along with the presentation of some prototype cold flow testing results which served as a successful proof of concept.
Description
| Type of resource | text |
|---|---|
| Form | electronic; electronic resource; remote |
| Extent | 1 online resource. |
| Publication date | 2014 |
| Issuance | monographic |
| Language | English |
Creators/Contributors
| Associated with | Waxman, Benjamin S |
|---|---|
| Associated with | Stanford University, Department of Aeronautics and Astronautics. |
| Primary advisor | Cantwell, Brian |
| Thesis advisor | Cantwell, Brian |
| Thesis advisor | Alonso, Juan José, 1968- |
| Thesis advisor | Zilliac, Gregory G |
| Advisor | Alonso, Juan José, 1968- |
| Advisor | Zilliac, Gregory G |
Subjects
| Genre | Theses |
|---|
Bibliographic information
| Statement of responsibility | Benjamin S. Waxman. |
|---|---|
| Note | Submitted to the Department of Aeronautics and Astronautics. |
| Thesis | Thesis (Ph.D.)--Stanford University, 2014. |
| Location | electronic resource |
Access conditions
- Copyright
- © 2014 by Benjamin Stephen Waxman
- License
- This work is licensed under a Creative Commons Attribution Non Commercial 3.0 Unported license (CC BY-NC).
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