I NASA - CR-54833 SRI Pr0jM-t No. 5075 4, j ,, FINAL REPORT ! EVALUATION OF OZONE DIFLUORIDE AS A LIQUID PROPELLANT ADDITIVE. BY A. B. AMSTER R. J. KlER E. L. CAPENER M. McCORMlCK L. A. DICKINSON J. NEFF Prepared for NATIONAL AERONAUTICS AND SPACE ADMINISTRATION December 1 1966 CONTRACT NAS3-4187 Technical Management NASA Lewis Research Center Cleveland, Ohio Erwin A. Edelman Polymer & Propulsion Sciences Division ICIILlI-Cl~h -I-cI-.-c,, mL.cI3-I.l- 3 1 HlYrUKLJ KC3CHKLn 11’43 I I I U I t 333 Ravenswood Avenue Menlo Park, California -94025 PRECEDING PAGE BLANK NOT FILMED. EVALUATION OF OZONE DIFLUORIDE AS A LIQUID PROPELLANT ADDITIVE A. B. Amster R. J. Kier E. L. Capener M. McCormick L. A. Dickinson J. A. Neff ABSTRACT The hypergolicity of OnF2 at the saturation point in liquid oxygen Tith liquid and gaseous hydrogen is critically dependent upon the tem- perature of the combustion chamber. When this temperature is above 150°K, pressure, total injected weight ratio of oxidizer to fuel and in- jector type were the primary variables of importance. I The preparation, mixing, and use of solutions of O3F2 in liquid oxygen is reviewed. The storage life of these solutions at 90°K is not expected to exceed one month. iii TABLE OF CONTENTS ABS"llJlCT .............................iii LIST OF ILLUSTRATIONS ....................... vii LIST OF TABLES ........................... ix SYMBOIS .............................. xi I SUMMARY ......................... ... 1 I1 INTRODUCTION ...................... ... 3 I I I EXPERIMENTAL APPROACH ..................... 5 A . General Dsscription .................... 5 B . Instrumentation ...................... 6 C . Rocket Combustor Design Factors .............. 9 IV PREPARATION OF 03F2 AND SOLUTIONS WITH LIQUID OXYGEN . ... 13 A . Preparation of 03F, .................... 13 B . Preparation of LOX/O3F2 Solutions .......... ... 13 V EXPERIMENTAL PROSRAM .................. ... 15 A . Hypergolicity Between 20 and 77'K ............. 15 B . Ignition Tests Using Combustor .............. 18 1. Experimental Procedure ................ 18 2 . Data Reduction .................... 21 3 . Experimental Results ................. 24 a . Engine Environmental Iemperature Variations . ... 24 b . Injector Influence ................ 28 c . Thrust kvel Chaages ............ ... 30 d . Influence of L* ...............-*a 32 e Altitude Ignitions with the Vortex Injector . **- 32 f . O,F, Concentration Level .......... ... 33 g . Stlidy of Gaseous Hydrogen Feed ....... * * * 38 C . Discussion of Results ................ ... 40 1. Ignition and Combustion Aspects ............ 40 2 . Factors Relating to Use in Large Rocket Engines . ... 43 V TABLE OF CONTENTS (Concl'd.) VI GENERAL CONCLUSIONS ...................... 55 ACKNOWLJ3DGMENTS .......................... 55 APPENDIX A .. Occurrence of Combustion Instability ........A-1 APPENDIX B ..Rocket Use Factors .IXIX/O,F, ............B-1 APPENDIX C .. Instrumentation and System Controls .........C-1 APPENDIX D .. A Simple Colorimeter for Assay of 0, Fa in Liquid Oxygen ..................D-1 vi LIST OF ILLUSTRATIONS Fig . 1 Flow System ....................... 6 Fig . 2 Instrumented Engine Assembly ............... 8 Fig . 3 Multi-coaxial Injector Assembly ............. 9 Fig . 4 Triplet Injector Assembly ................ 10 Fig . 5 Impinging Vortex Injector Assembly ............ 11 Fig . 6 Mix Tank for O,F, .................... 14 Fig . 7 Vial Design for Hypergolicity Test with Liquid Hydrogen ..................... 16 Fig . 8 Temperature Distribution in LOX/O,F, Sphere at 7fK When Plunged Suddenly Into Liquid H, at 20°K ....... 17 Fig . 9 Double Dewar Assembly for Controlling Engine Start Temperature ....................... 19 Fig . 10 Kistler and Ion Gages Oscilloscope Traces Showing Ignition Outside Engine ................. 23 Fig . 11 Ignition Delays Inside Engine at Low Tempzrature and Low Flow Rates .................... 26 Fig . 12 Ignition Outside Engine at Low Temperature and High Flow Rates ................... 27 Fig . 13 Ignition Delays for Engine Environmental Temperatures from 190'K to 290'K ............. 29 Fig . 14 L* Variations with the Triplet Injector ......... 30 Fig . 15 Effect of Altitude Using Vortex Injector ......... 31 Fig . 16 Effect of 0, Fa Concentration on Ignition Delay ...................... 34 Fig . 17 Analysis of LOX/O, F2 Solution Delivered at the Rozket Engine ...................... 38 Fig . 18 Ignition Delays Obtained Using Gaseous Hydrogen ..... 39 Fig . A-1 Conbustion Instability Feedback to Injector Pressures. U>X/O,F, Liquid Hydrogen Engine ........A-1 Fig . B-1 (a) Apparatus for Synthesis of O,F, ...........B-3 (b) Apparatus for Synthesis of 0,F2 ...........B-3 (c) U-Tube O,F, Reactor .................B-3 (d) U-Tube O,F, Reactor .................B-4 (e) Multiple Reactor ...................B-4 rig . B-z (aj MX-Cooiea tiiass Piper;r;e uzwar r'iase ~ac~et.... .B- IZ (b) LOX-Cooled Funne 1 ..................B-12 (c) LOX-Cooled Pipette (Double Housekeeper Seal) .....B-12 vi i LIST OF ILLUSTRATIONS (Concl 'd . ) Fig . C-1 Basic Instrumentation Circuits ..............C-2 Fig . C-2 Thermocouple Construction ................C-6 Fig . C-3 Kistler Gage Mounts ...................C-6 Fig . C-4 Ion Probe Construction ..................C-7 Fig . C-5 Automatic Sequencer ...................C-7 Fig . D-1 Source and Dztector Assembly ...............D-2 Fig . D-2 Bridge Circuit ...................... D-3 Fig . D-3 Calibration Data For Colorimetric Analyzer ........D-5 viii LIST OF TABUS Table I Summary of Test Conditions .............. 20 Table I1 Test Series 2. .................... 45 Table I11 Test Series 3. .................... 46 Table IV Test Series 4. .................... 47 Table V Test Series 5. .................... 48 Table VI Test Series 6. .................... 49 Table VI1 Test Series 7. .................... 50 Table VI11 Test Series 8. .................... 51 Table IX Test Series 9. .................... 52 Table X Test Series 10 .................... 54 Table A-I Peak-to-Peak Amplitudes of Combustion Waves Generated with the Impinging Vortex Injector .....A-3 Table B-I Physical and Physicochemical Properties of Trioxygen Difluoride .................B-6 Table B-I1 Chemical Reactivity of 0,F2. .............B-7 Table B-I11 Open Cup Tests of 0.05% O,F, with Various Fuels. ...B-7 Table B-IV Cospatibility Tests of Engineering Materials with 0.05% 0,F2 (By Weight) Solution in Liquid Oxygen. .................... B-9 Table C-I Basic Instrumentation Data ..............C-3 Table D-I Calibration Data for Colorimetric Analyzer ......D-4 ix SYMBOLS V Velocity - ft/sec P Pressure - psi P Density - lbs/ft3 g Gravity - ft/sec" Z Height - ft S Are? - ft" - Pi (ef f 1 Effective injector pressure - Psi t Time - sec Instantaneous injector pressure - psi 'i W Flow rate, steady state - lbs/sec 0 W Equivalent flow rate during unsteady time interval - Ibs/sec At Ignition delay - milliseconds 0 0 Total integrated flow of oxidizer up to ignition time - lbs 0 F Total integrated flow of fuel up to ignition time - lbs T Absolute temperature - Degrees K L* Chamber volume/nozzle throat area - ft Subscripts: 1 Station 1 2 Station 2 3 Station 3 0 Oxidizer i Injector f Fue 1 L Flow line T Tank C Engine chamber R-L: ---L .. ..u niuuiciu L xi The over-all results of our investigation indicate that hypergolicity of O,F, is associated with its thermal decomposition, which results in the formation of reactive radicals. Rapid ignition is associated with relatively warm combustion chambers (above 150'K); it also depends on the ambient pressure. A significant dependence of ignition delay on the time-averaged weight ratio of oxidant to fuel was found for all injector 00 types; however, at comparable O/F ratios, delay was dependent also on the injector type. A number of problems associated with instrumentation and 0,F2 han- dling were encountered; some of these remain unresolved such as quantita- tive calibration for transient , multi-phase flow through orifices, and analytical problems associated with the use of a highly reactive and thermally unstable additive. Significant advances were made experimentally in determining critical ignition parameters such as transient propellant mass flow rate and the time and place of ignition. It is believed that the experimental results provide the engineering background necessary for any future consideration of scale-up to larger rocket engines. While some progress was made in the field analysis of LOX/03F2 solutions, a good method for characterizing LOX/03F2 suspensions was not developed. This study has shown that rather critical requirements must be met if satisfac- tory ignition is to be achieved; hence O,F, is not a simple panacea for eliminating ignition or combustion problems in rocket systems that use liquid oxygen. While the practicality of using O,F, in space propulsion cannot be unequivocally stated without a detailed engineering systems study, some critical conditions exist. These may be summarized as follows: 1. Liquid oxygen/O,F, solution must be delivered to the injector face if the O,F, is to be effective as an ignition parameter. 2. Tile comuustion chamber should be at a temperature above 150°K. 1 3. The storage life of prepared mX/0,F2 solution at 90°K is not expected to exceed one month; the storage life of solutions at temperatures above 90°K has not been determined. 4. The relatively long ignition delays of 5 to 50 milliseconds observed-may preclude the use of mX/O3F2 in small impulse bit devices. It has been suggested that O3F2, by modifying