The objective of the current work is to experimentally characterize the flowfield associated with an uni-element shear coaxial injector burning liquid oxygen/gaseous hydrogen (LOX/GH2) propellants. These experiments were carried out in an optically-accessible rocket chamber operating at a high pressure (approximately 400 psia). Quantitative measurements of drop size and velocity were obtained along with qualitative measurements of the disintegrating jet

Foust, M. J.

Moser, M. D.

Pal, S.

Ryan, H. M.

Santoro, R. J.

English

NASA Technical Reports Server

N95- 23642
Flowfield Characterization in a LOX/GH z Propellant Rocket
S. Pal, M. D. Moser, H. M. Ryan, M. J. Foust and R. J. Santoro
Propulsion Engineering Research Center
and
Department of Mechanical Engineering
The Pennsylvania State University
University Park, PA 16802
f
/'
Statement of Problem
There is a critical shortage of data pertaining to the flowfield characteristics in a liquid propellant
rocket chamber for hot-fire conditions. For a liquid oxygen (LOX)/gaseous hydrogen (GHz) propellant
combination, either shear or swirl coaxial injectors are typically used to atomize the liquid propellant into
drops. Understanding the atomization process under hot-fire conditions represents the first step in
understanding the subsequent processes of vaporization, mixing and combustion. The flowfield here is
two-phase and therefore experiments that detail the intact liquid jet, drop size and velocity, and
combustion length are necessary for understanding the physics of the problem.
Objective of Work
The objective of the current work is to experimentally characterize the flowfield associated with
an uni-element shear coaxial injector burning LOX/GH 2 propellants. These experiments were carried out
in an optically-accessible rocket chamber operating at a high pressure (--400 psia). Quantitative
measurements of drop size and velocity were obtained along with qualitative measurements of the
disintegrating jet.
Approach
The experiments were conducted at the Cryogenic Combustion Laboratory at Penn State
University. This laboratory provides the capability for firing both gaseous and liquid propellant sub-scale
rocket engines. A modular rocket chamber which provides extensive optical access was designed for the
experiments. The cross-section of the rocket is 50.8 mm (2 in.) square and the chamber length which
can be easily varied was 248 mm (9.75 in.). The fiowfield downstream of a shear coaxial injector was
characterized using laser-based diagnostic techniques. The inner diameter of the LOX post was 3.43 mm
(0.135 in.) and the post was recessed 3.78 mm (0.15 in.). The inner diameter of the fuel annulus was
4.19 mm (0.165 in.) and the outer diameter was 7.11 mm (0.28 in.). The nominal LOX and GH 2
flowrates were 0.11 kg/s (0.25 Ibm/s) and 0.021 kg/s (0.047 Ibm/s) respectively, thus resulting in a
nominal O/F ratio of 5.3:1. These flow rates, coupled with the nozzle dimensions yielded a chamber
pressure of 2.74 MPa (=400 psia). The duration of a test run was four seconds.
A Phase Doppler Particle Analyzer (PDPA) was used to measure LOX drop size and velocity at
various radial locations for an axial position 63.5 mm (2.5 in.) downstream of the injector face. The
measured Sauter Mean diameter (SMD) ranged from 110 #m at the centerline to about 60 #m at a
9.5 mm (0.375 in.) radial location. At greater radial locations, no drops were observed. The results
indicate that under hot-fire conditions, the drops formed from the shear coaxial injector are confined to
a narrow circumferential region.
Conclusions
These experiments represent the first time that drop sizes have been measured under combusting
conditions for cryogenic propellants. These results are, in general, encouraging with respect to
applications of laser-based diagnostics to LOX/GH2 uni-element rocket studies. A comprehensive
mapping of the flowfield will need to be completed to gain a thorough understanding of the physics of
this complex problem.
613
https://ntrs.nasa.gov/search.jsp?R=19950017222 2020-06-16T08:47:06+00:00Z
Flowfield Characterization in a Liquid
Oxygen/Gaseous Hydrogen Propellant
Rocket
S. Pal, M. D. Moser, H. M. Ryan, M. J. Foust and
R. J. Santoro
The Propulsion Engineering Research Center
The Pennsylvania State University
University Park, PA 16802
llth Workshop for CFD Applications
in Rocket Propulsion
April 20-22, 1993
NASA Marshall Space Flight Center
Huntsville, Alabama
614
PRESENTATION OUTLINE
Motivation
Background
Objective
Experimental
- Facility
- Rocket chamber
- Diagnostics
Results
Summary
615
MOTIVATION
To obtain fundamental data under well characterized
conditions
- For gaseous propellant rockets
- For liquid propellant rockets
- For various injector types
Fundamental data would form the basis for
- Empirical correlation validation
- CFD code validation
616
BACKGROUND
Atomization models typically are:
- Anchored to cold flow experimental results
We and Re of cold flow experiments differ by an
order of magnitude from actual rocket conditions
Results have to be extrapolated
- Predicted from analytical models based on linear
stability theory
!
Drop size data for hot-fire conditions would:
- Validate atomization models
- Validate methodology for extending cold flow data to
hot flow conditions
- Develop hot-flow correlations for direct use in
combustion models
617
OBJECTIVE
To characterize the flowfield downstream of a shear
coaxial injector using LOX/GH 2 propellants under
combusting conditions
- Drop size and velocity measurements using Phase
Doppler Interferometry
- Laser sheet visualizations of near breakup region
618
FACILITY CAPABILITIES
Propulsion Engineering Research Center
Cryogenic Combustion Laboratory
Propellants:
Gaseous Hydrogen
Gaseous Methane
Gaseous Oxygen
Liquid Oxygen
Flow rates (maximum):
Gaseous oxygen: 0.1 lbm/s
Liquid oxygen: 1.0 lbm/s
Gaseous hydrogen: 0.25 lbm/s
Typical mixture ratios: 4 - 8
Maximum chamber operating pressure: 1000 psi
Typical test time: 1 - 5 s
Longer tests subject to hardware cooling and gas
supply specifications
619
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TEST CONDITIONS
Chamber pressure 2.67 Mpa (387 psia)
LOX flowrate 0.112 kg/s (0.245 Ibm/s)
GH2 flowrate 0.021 kg/s (0.045 Ibm/s)
Mixture ratio (O/F) 5.4
LOX velocity
GH 2 velocity
13.5 n!/s (44.1 ft/s)
381 m/s (1250 ft/s)
Momentum ratio (F/O) 5.22
Velocity ratio (F/O) 28.3
Rd 5.03x10 s
Wd 2.06x10 s
_Re = p_U_ d//h
2We -- pg(Ug-Ui)2d/_r
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SUMMARY
Measured liquid oxygen drop size and velocity in
combusting environment
- Intact core extends well beyond the injector
- Drops confined to narrow region
Correlations based on cold flow data inadequate for
predicting drop size in LOX/GH 2 combusting flow
Compared drop measurements between cold flow and
combusting conditions for similar liquid and gas
flowrates
- Re I and We e differ by an order of magnitude
- Mean drop size larger for hot fire conditions
643
ACKNOWLEDGEMENT
Funding by NASA Marshall Space Flight Center, contract
NAS8-38862 and the Penn State NASA Propulsion
Engineering Research Center, Contract NAGW 1356
supplement 2
644


Flowfield Characterization in a LOX/GH2 Propellant Rocket

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