effect of variation of chamber geometry on the performance of a small scale bipropellant thruster
- Paper number
IAC-07-C4.3.08
- Author
Mr. Arun Kumar P , Liquid Propulsion System centre, India
- Coauthor
Mr. S. Venkateswaran, Liquid Propulsion System centre, India
- Coauthor
Mr. S. Kothandan, Liquid Propulsion System centre, India
- Coauthor
Mr. Ajith B, Liquid Propulsion System centre, India
- Year
2007
- Abstract
Extended Abstract:
LPSC is developing a 5N bipropellant thruster to meet the future requirements of maneuvers for different torque arm spacecrafts. As a part of design efforts, a study is undertaken to evaluate by experiment the effect of variation in combustion chamber design parameters on the performance and this paper presents the details. Propellant combination is hypergolic Nitrogen Tetroxide (NTO) / Monomethyl Hydrazine (MMH).
The thrust chamber is the combustion device where the liquid propellants are injected, atomized, vaporized, mixed and burnt to form hot gaseous reaction products, which in turn are accelerated and ejected at high velocity. For the thrust chamber design, the starting point is to finalize the combustion chamber geometry followed by nozzle design.
Even though the basic atomization is described in many of the literature, the combustion mechanism of hypergolic propellants is different from non hypergolic of which few published literatures are available on this subject. The total combustion process, from injection of reactants until completion of the chemical reactions and conversions of the products into hot gases, requires finite amounts of time and volume, as expressed by the characteristic length , L*. The value of the factor is significantly greater than the linear length between injector face and the throat plane. The contraction ratio, which is defined as the major cross sectional area of the combustor (typically in the cylindrical position near the injector face) divided by the throat area, is basically related to the mean Mach number of the combustion gas.
Typically, large scale engines are constructed with a low contraction ratio and a comparatively long length , and smaller chambers employ a large contraction ratio with a shorter length, while still providing sufficient L* for adequate vaporization and combustion dwell time. In general, combustion process is not scalable and optimum combustion chamber geometry is arrived at by experimentally evaluating the performance by changing L* and contraction ratio.
Experimental injector and thrust chamber
Injector is a coaxial swirl single element and is made of Ti6Al4V. In a swirl injector, the liquid is fed into the swirl chamber through tangential inlet channel. At the injector nozzle exit, the jet will be converted into a conical film, which is further broken into drops. This inherent characteristic of the injector results in an unintentional film cooling of the thrust chamber.
The thrust chamber is sea level version and is made of Columbium alloy, silicide coated and radiation cooled. The valves used are of the solenoid type with sliding plunger. The injector assembly is connected to the thrust chamber by a screw joint.
Five sea level version thrust chambers were realized by varying contraction ratio and L* as given in table 1. By varying these design parameters, the distance of the throat from the injector face (Lt) is changed.
No Chmb. L*(cm) Lt(mm) Cntn.ratio
001 SL-01. 80.000 42.70000 23.04
002 SL-02. 90.000 47.15000 23.04
003 SL-03. 100.00 51.50000 23.04
004 SL-04. 90.000 52.56000 19.36
005 SL-05. 100.00 57.63000 19.36
Table 1 Details of combustion chambers realizedExperimental Results
Performance evaluation was carried out by doing sea level continuous firing for 100 s for the five chambers with a common injector. Chamber pressure (Pc) was maintained at 6.65 ±0.35 bar and Mixture ratio (MR) 1.60 ±0.05.
In the test facility, mass flow meter with an accuracy of ±0.5
All the tests were carried out at an injection pressure of 16.5 bar (a) for both oxidizer and fuel and the test results are summarized in table 2. Performance parameter evaluated is C*eff, which is computed by C*eff = Pc(×)At(×)100(×)g0/(mt(×)C*theor)
where At = Throat area mt = Total propellant mass flow rate C*theor= Theoretical C* which is 1737m/s for NTO/MMH at 1.60 MR and 6.65 bar (a) go=9.81m/s2
Chmb M.R. Pcbar C*eff Thrt temp(degC)
SL-01 1.630 6.300 90.100 960
SL-02 1.580 7.000 93.200 1010
SL-03 1.600 6.300 94.300 1060
SL-04 1.580 6.300 97.100 1230
SL-05 1.550 6.610 96.400 1150
Table 2 Test summaryEffect of L* on performance:
With reference to the tables 1 and 2, for a fixed contraction ratio, C*eff as well as throat temperature increases with L*. This is due to increased Lt resulting in better residence time of product of combustion.
Effect of contraction ratio on performance
With reduced contraction ratio, C*eff improves and throat temperature increases as indicated by the test summary. This is again due to the increased Lt for lower contraction ratio. This shows the physical constrain of a small scale thrust chamber which limits the achievable performance as the contraction ratio will be larger compared to that of a large scale thrust chamber.
The effect of chamber length (Lt) on the performance can be evaluated by combining L* variation and Contraction ratio variation datas. C*eff and throat temperature increases with Lt. An interesting observation is that the sensitivity of the performance (C*eff) to the chamber length (Lt) was greater for Lt between 42 to 51.5 mm compared to those between 51.5 to 57mm. The throat wall temperature depended heavily on chamber length which not only controls the C*eff but also the effectiveness of film cooling at the throat.
Concluding remarks:
Experiments were carried out by hot firing for 100 sec, a coaxial swirl injector with five seal level version chambers with varying design parameters. The results are summarized as follows:
1.C*eff is a function of chamber length (Lt) irrespective of L* and contraction ratio.
2.Sensitivity of C*eff to chamber length is greater for Lt between 42 to 51.5mm compared to those beyond 51.5mm.
3.Throat temperature also depends on chamber length as it controls the C*eff and film cooling effectiveness.
4.A compromise between performance and throat wall temperature determines the optimum chamber length of a film cooled small range thruster.- Abstract document
- Manuscript document
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