Metallized gelled propellants - Oxygen/RP-1/aluminum rocket heat transfer and combustion measurements

Transcription

1 Copyright 1996, American Institute of Aeronautics and Astronautics, Inc. AIAA Meeting Papers on Disc, July 1996 A , AIAA Paper Metallized gelled propellants - Oxygen/RP-1/aluminum rocket heat transfer and combustion measurements Bryan Palaszewski NASA, Lewis Research Center, Cleveland, OH James S. Zakany NYMA, Inc., Brook Park, OH AIAA, ASME, SAE, and ASEE, Joint Propulsion Conference and Exhibit, 32nd, Lake Buena Vista, FL, July 1-3, 1996 A series of rocket engine heat transfer experiments using metallized gelled liquid propellants was conducted, using a small lbf thrust engine composed of a modular injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-, 5-, and 55-wt pct loadings of Al particles. Gaseous oxygen was used as the oxidizer. Three different injectors were used during the testing, one for the baseline O2/RP1 tests and two for the gelled and metallized gelled fuel firings. Heat transfer measurements were made with a rocket engine calorimeter chamber and nozzle with a total of 31 cooling channels. Each channel used a water flow to carry heat away from the chamber and the attached thermocouples and flow meters allowed heat flux estimates at each of the 31 stations. The rocket engine characteristic velocity efficiency for the RP-1 fuel was in the percent range, while the gelled 0-wt pct RP-1 and the 5-wt pct RP-1 exhibited an efficiency range of percent and percent, respectively. The 55 wt pct RP-1 fuel delivered a percent efficiency. (Author) Page 1

2 METALLIZED GELLED PROPELLANTS: OXYGEN/RP-1 /ALUMINUM ROCKET HEAT TRANSFER AND COMBUSTION MEASUREMENTS Abstract A series of rocket engine heat transfer experiments using metallized gelled liquid propellants Bryan Palaszewski NASA Lewis Research Center Cleveland, OH James S. Zakany NYMA, Inc. Lewis Research Center Group Brook Park, OH 44142

3 Ibf MMH pound-force Monomethyl Hydrazine

4 an issue, due to the small size of the engine and the relatively large purge flow rates needed to make the engine firing repeatable. Questions also arose

5 were sought and obtained. Testing was therefore conducted with O 2 /RP-1/A1 propellants using gelled RP-1 with aluminum particles. During

6 Combustion Testing Description

7 inside diameter,

8 and oxidizer manifolds,

9 where: CF = Thrust coefficient at nozzle exit (using Ref. 22) PC = Chamber Pressure MPL

10 Figure

11 wt% cases

12 oxide coating

13 estimates from

14 gel only) was evident in the fuel dome once it

15 realistic flow conditions, reduced influence of igniter purges, and allow researchers to gather more engine data in a more-representative highthrust rocket environment. Acknowledgements We'd like to thank NASA Headquarters, Office

16 14) Palaszewski, B., "Metallized Propellants for the Human Exploration of Mars," NASATP-3062, November Also, AIAA Journal of Propulsion and Power, Nov.-Dec

17 Table

18 n Table H (continued) Calorimeter Engine Geometry Data Nozzle: Axial Location from injector face (cm) where: n is the axial station (starting at the injector face) Table EEL. Typical Metallized Gelled Propellant Engine Flow Rates

19 200 +? 180.

20 \ Injector Spool piece or combustion chamber Nozzle BAP VMM Coolant channel inlet / outlet tubes (22 each for chamber. 9 each for nozzle) Figure 3. Simplified Diagram of Engine Components and Configuration (T O % wt% RP-1/AI: Calorimeter « ii I 0.41 E co O J o.o o o B 1 2 3

21 * Theoretical Vacuum Isp A Test Data: Upper

22 : Q. V) g 60.1 o> 'o i RP-1 0% 5% 55% Propeilant type Figure 7. Summary of Isp Efficiency Ranges for all Propellants: At Peak Isp ^ 200 o Q. RP-1: Runs ^ 1501 S ^T 1001 u> a. o 501 «^o V) Main Combustor

23 a a P, _, v> Specific impulse efficiency ro 4^ CD c> o o o o o O 1j ^ i... VO O co )-* Cstar efficiency ro 4s» o> co to to 1 ss* 1 1? ** H»" O N 3.j, o> a 5* 0 0 K» 3 M ' o- c (0 o -i w. *** Q n 4^. Oi 1 3 rd ^ *1 ^w 31 33? S 5 > I C ^ 3 "^ Crt (A ^ O I? t * 1" P

24 Cstar efficiency Specific impulse (Ibf-s/Ibm, PC) \v to b o IO 4^ O 00 o O O O O O O [ a 1I * *» oo to at o O O O O O O O 1 B S jf o *r? 1t k g -*- 0) 3 o O cr c»0 (0 o 5* ta. t n El,., t e OL 0 B ft T*H 0 1HJ <g 3} 31 c 3 (O O 1 (O en OS 6J 1 a I9^ 1 s< 1 OI

25 Specific impulse (Ibf-s/Ibm, Specific impulse efficiency

26 iuir 5 wt% RP-1/AI: Runs c 0) "5 60 o> L. CO to 40-0 BG FT ET B3 «a m QTI i m 20 Figure n Main Combustor O/F

27 Cstar efficiency Specific impulse (Ibf-s/Ibm,

28 p w 13 O 5 s s I oi u Heat flux (MW /m2) -^ to co en o 09 CL 9

29 2.0 RP-1 CM.,

30 Runs 873 (RP-1) and 986 (5 wt%),-s 7 " CM E 61 RP-1 O RP-1/AI: 5 wt% O $ 41 0) X 3' 2" O O QB "

31

32 Injector Layer