Held on Saturday,March 25, 1995 at the HALO Rocket Motor Test Facility
The following text was taken, with permission, from an article by Tim Pickens, which was published in the May-June 1995 issue of the Southeastern Space Supporter, newsletter of HAL5.
It was Saturday, March 25, and the day of reckoning had finally arrived. Much work had been performed preparing for this day. The test facility was to be christened with an asphalt and nitrous-oxide (N2O) hybrid rocket motor (in the middle of the stand test) capable of achieving a total impulse of 2000 lb-sec (200 pounds of thrust for 10 seconds), which we estimate could loft a balloon-launched rocket into space. The large gauge would monitor the chamber pressure inside the motor, while the smaller one would monitor the pressure inside the oxidizer tank (protruding above the test stand). Electronic monitoring was also to be performed using a circuit board (inside the metal box) to relay data back to a computer in the HALO control room (that is, the barn of Herman Pickens).
There were many unknown factors. Although asphalt is a common roof coating (and road material), it was difficult obtaining an exact chemical formula and other pertinent design information. Asphalt offers an affordable way for us to develop, produce, and test our instrumentation, flow system, ignition system, and other motor-related components. HAL5 is very fortunate to have talented people as members of Project HALO to allow us to do 100 percent in-house design, construction, and testing.
The major theoretical motor design aspects were accomplished by Steve Mustakis, who is a propulsion major about to earn his Bachelors degree from UAH. He has been very instrumental in all aspects of design and construction.
Our hybrid motor casing consisted of a 2.5-inch diameter, 23-inch long steel pipe. Each end (injector and nozzle) was threaded for screw-on caps, which allowed for quick assembly and disassembly. The fuel grain consisted of a 15-inch long cylinder of asphalt with a 1-inch core (hole) down the center. The separate motor nozzle was a 3-inch long Delavel design carved from solid graphite (in Tim Pickens garage). Separating the asphalt fuel from the nozzle was a 4-inch, post-mixing combustion chamber.
With the initial fuel grain weighing in at 1.6 lbs, Steve calculated that about 10 lbs of N2O would be required to achieve the best rocket performance (based on the optimum fuel-to-oxidizer ratio). To achieve the desired 200 lbs of thrust for 10 seconds, we needed an average oxidizer flow rate of 1.0 lb/sec. We first tried to design an oxidizer injector with a single central port (hole), but calculations showed that the oxidizer would not expand properly and might cause fuel regression problems near the injector. We chose to go instead with a multi-port injector design, which proved to have very good expansion properties. A cold flow test of our test injector showed that we could achieve the desired flow rate.
At the HALO Rocket Motor Test Stand, our oxidizer flow system was oriented vertically above the motor, a flight-ready configuration. This was done to allow us to gather realistic flow data. A pressure tap allowed us to examine the N2O blow-down characteristics during the test. The flow system was to be driven by the vapor pressure of the N2O inside the oxidizer tank; no pumps were to be used. This was one of the factors in selecting N2O, although there are some performance penalties. N2O is very safe and easy to work with, and can survive the projected two-hour balloon trip to the launch altitude.
The oxidizer valve we used (for its simplistic design) was a ball valve that was spring loaded to fully open. A string was wrapped around a pulley attached to the valve to hold the valve in a closed position. A squib was placed near the string. Once the squib ignited, the string would break and the spring would pull the valve open. Many successful ground tests of the valve were performed. A high altitude test of the valve never occurred due to a cold battery aboard the balloon.
Our ignitor would be a very critical area of our hybrid test program. Faulty ignitors are the most probable cause of past rockoon failures. We needed an ignitor with low power requirements because we were attempting to keep the future rocket weight at a minimum, and batteries are relatively heavy. The ignitor would be required to burn for a minimum of three seconds, generating enough heat to vaporize the surface of the fuel grain near the injector and ensure a good ignition when the oxidizer started flowing past. We needed a considerable amount of heat (about 570 deg-F) in order to disassociate the oxygen from the N2O.
Our ignitor design of a cylindrical wire mesh wrapped in Thermalite (described in the Mar-Apr issue) achieves all of these requirements. The squibs which ignite the Thermalite can be fired using a common watch battery. Our ignitor burns for about five seconds, then is totally consumed by fire once the motor itself ignites.
All data was to be recorded on a Macintosh computer inside the control room. HALO member Larry Larsen had set up a very nice system (using Lab View software) that would make data acquisition a state of the art affair. The command controls would be manual by toggling a series of switches. The switches were connected to a relay card, which would allow us (eventually) to run the controls via the computer.
The motor had sensors for monitoring thrust, starting and flow pressure, and chamber pressure. HALO member Gene Hornbuckle built amplifiers to increase the millivolt output of the sensors up to 10 Volts. Gene packaged the amplifiers and relays onto a single board which sits in a metal box next to the test stand when in use; and can be easily removed after. Gene, Larry, and Steve worked long hours refining and calibrating the equipment.
The ignitor was slid into the motor, then the motor was strapped to the test stand and connected to the oxidizer tank via the oxidizer valve mechanism. We then sounded an alarm siren briefly to clear the immediate area so that we could begin loading the oxidizer tank. This was a safety issue because we wanted to test a lightweight tank that we could later use for the flight vehicle. We wanted as much data as we could obtain on all aspects of our designs.
The N2O supply bottle was kept safely behind a concrete wall at one end of the test stand. Two people (Steve and myself) stood next to the bottle and filled the tank from behind the wall. (We plan to eventually go with a much safer remote loading capability.) The bottle sat on a scale so that we could monitor the change in weight as we filled the oxidizer tank. Knowing the internal volume, pressure, and mass of the oxidizer in the tank, we could compute the tank ullage (the amount of relatively useless gaseous N2O floating above the thrust-enabling liquid N2O below). After filling the tank, we removed ourselves to the safety of the control room.
Our rocket motor test system is completely safe until the oxidizer tank is loaded. Even then, two events must occur to start the test. First, the ignitor must be triggered; second, the oxidizer valve must be opened. We even had a backup for the oxidizer valve; a long string between the valve and the control room, which could be used to yank open the valve if the squib failed.
After a second and final siren warning, we counted down five seconds, then threw the switch. With a loud snap, the ignitor squib fired and started the ignition process, heating the asphalt and filling the combustion chamber with fuel vapor. Two seconds later, we fired the valve squib, which successfully opened the oxidizer valve and allowed the N2O to enter the chamber. Whoosh! Motor ignition! All eyes were focused on this historic moment for Project HALO.
The test looked great! This test seemed to burn forever, although the video recording proved that the actual burn time was only eight seconds. Motor ignition was very good and thrust quickly climbed to a peak. Chamber pressure went as high as 700 psi, 300 more than our design requirement (we would later correct this). It was evident from our pressure gases that there was coupling (chugging) taking place between our chamber pressure and our blowdown pressure.
Unfortunately, our data acquisition system still had a few bugs, and our data was very noisy. (We later corrected this by replacing the power supply in the metal box.) We believe we achieved a thrust of about 125 lbs at a specific impulse (Isp) of about 210 seconds. All data indicated that we were with 25% of our theoretical predictions.
The only casualty of the test was a wooden flame deflector, which looked like burnt toast after the test. The motor flame burned a hole right through the deflector and gouged out a small crater in the concrete pad! All in all, however, we were very pleased with our first test firing.
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This file was last modified on Saturday, 15-Apr-2017 13:19:39 EDT