Development of 2500 N Class CAMUI Type Hybrid Rocket for Winged Flight Experiments

Paper number

IAC-10.C4.2.9

Author

Prof. Harunori Nagata, Hokkaido University, Japan

Coauthor

Dr. Masashi Wakita, Japan

Coauthor

Prof. Tsuyoshi Totani, Hokkaido University, Japan

Coauthor

Mr. Tsutomu Uematsu, Japan

Coauthor

Prof. Koichi Yonemoto, Kyushu Institute of Technology, Japan

Year

2010

Abstract

The authors have developed CAMUI type hybrid rockets as an explosive-free small launch system. A main purpose is to reduce experimental expense of small scale flight experiments by cutting back on safety management cost. CAMUI comes from an abbreviation of a new combustion method of 'cascaded multistage impinging-jet'. By separating conventional cylinder-shape solid fuel grain with a central port into multiple cylinder blocks, the combustion gas repeatedly collides with forward end faces of fuel blocks to accelerate the heat transfer to the fuel. A 2500 N class CAMUI motor is under development for a small scale winged flight test bed. 

The motor uses polyethylene and liquid oxygen (LOX) as propellants with the total amount of 5 kg, excluding a residual fuel sliver. Target thrust is above 2500 N at peak and 2000 to 2500 N on average. Desired specific impulse is 250 sec, meaning that the available impulse from the 5 kg propellant is 12.3 kNs. The motor of 20 kg in total weight consists of three high pressure helium tanks of 5 liters in total capacity, a 3.5 liter LOX tank, and a thrust chamber. The fuel grain consisting of eight fuel blocks is 410 mm in length, 100 mm in diameter, and 1.9 kg in weight.

Static firing tests provided successful thrust performance of the motor. Test data includes combustion chamber pressure, LOX flow rate, and thrust. After a firing test, retrieve a remaining fuel sliver to measure the weight. From histories of LOX flow rate and chamber pressure,  temporal variation of fuel gasification rates was estimated by using a reconstruction technique. The peak and mean thrust were above 2,600 N and about 2300 N, respectively. A fuel sliver retrieved after a firing test weighed  0.45 kg, meaning that the fuel consumption during the firing was 1.45 kg. Adding LOX weight of 3.85 kg, the total propellant consumption was 5.30 kg. Integration of thrust history gave the total impulse to be 13.2 kNs, exceeding the target value of 12.3 kNs. Dividing the total impulse by the total propellant consumption gave mean specific impulse to be 254 sec, achieving the target value of 250 sec. This motor is able to launch a winged vehicle of 50 kg to about 1.8 km apogee altitude. The launch experiment is slated for this May or June.

Abstract document

IAC-10.C4.2.9.brief.pdf

Manuscript document

IAC-10.C4.2.9.pdf (🔒 authorized access only).

To get the manuscript, please contact IAF Secretariat.