Lotus Dev 2

After a successful test campaign of Lotus Dev 1 (LD1), BURPG is introducing the latest engine to the Lotus series, Lotus Dev 2 (LD2). LD2 will power BURPG’s Starscraper suborbital launch vehicle to 100km in the summer of 2017.

The pressure-fed engine produces 2,500 lbs of thrust with 233 sec Isp SL. LD2 features an increased chamber pressure and L* for boosted engine performance.  A single LD2 will replace the clustered LD1 configuration for Starscraper’s spaceshot.

Lotus Dev 2

To support development of LD2, BURPG will employ Iron Lotus - a steel heat-sink engine with the internal chamber geometry of LD2. Iron Lotus will allow the team to complete ignition testing prior to putting LD2 on the stand; getting BURPG to the test stand faster and with less hardware risk.  Ignition data collected will be used to push LD2’s aggressive test campaign.

Iron Lotus

In order to obtain reliable and repeatable ignition, BURPG has turned to Triton Space Technologies’ TS-160Y pyro actuated poppet valves for the engine's main valves. The TS-160Y provides up to 3,000 psi working pressure and is coupled with high flow capacity. It features 1" male AN connections, a liquid CV of 13, and an open response time of < 75 ms - faster than typical pneumatically actuated ball valves. BURPG is lucky to be working closely with Triton Space Technologies in support of the Starscraper program. Triton Space Technologies' technical and financial support will go a long way as BURPG pushes the envelope of amateur aerospace. Read more about Triton Space Technologies (www.triton-space.com) and the TS-160Y valves here: www.triton-space.com/ts-160y.

TS-160Y Valves

In parallel with LD2 development, vehicle design continues. Starscraper subsystems have undergone both internal and external CDR's for the new intertank, fuel tank, feed system, and TVC system. A full vehicle CDR is scheduled for the month of December as we aim to complete flight paperwork before the end of the year.

BURPG is excited for the months ahead as Starscraper build ramps up and engine testing progresses. If you would like to support our team, please email us at terrierrocket@gmail.com.

Year In Review

The 2015-2016 school year saw BURPG kicking off development of a new engine, Lotus. Replacing the hybrid Mk. V on the Starscraper rocket, Lotus is BURPG’s first crack at liquid engine development, and has proven to be an incredible learning experience and significant step in the journey to space. As the group takes a semi-hiatus from rocket development for the summer, it’s time to look back on the progress with Lotus, Starscraper, and where things are going next.

Lotus liquid engine on the test stand

Lotus liquid engine on the test stand

One of the major reasons for BURPG’s transition to liquid engines is the faster testing pace they offer. Lotus testing began in December 2015, after a marathon build and prep process during the fall semester. In just the spring semester alone, BURPG executed more engine tests than ever before accomplished in an entire year. Not only that, but Lotus testing requirements birthed a totally new ground support system, software, and upgraded electronics that were refined and matured through the test campaign. BURPG now has a very capable, flexible and reliable test infrastructure.




BURPG executed 26 tests of Lotus during the spring semester. Six of these tests were aborted before or shortly after engine ignition due to the detection of faults in either the ground system or the engine itself. While these tests may not have provided data about engine performance, they still yielded data used to improve the engine or ground support system. There were 10 cold flows of Lotus.

Cold flow test 

Cold flow test 

Hot fire test

Hot fire test

There were 10 hot fire attempts of Lotus, 8 ignited, with all tests being 2 second duration. These tests focused on ignition and establishing stable combustion. A brief breakdown of the tests are as follows:

Testing Summary:



Equipment Notes

December 5, 2015

4 oxidizer cold flows

3 fuel cold flows

-Rev 1 fluids

February 20, 2016

1 fuel cold flow

3 hot fires

-Rev 1.1 fluids: reduced propellant line lengths


April 3, 2016

1 hot fire

3 aborted tests

-Rev 2 fluids: panel mounted pressurization system, further optimized propellant lines

-Corrected injector fuel port sizing

April 17, 2016

1 fuel cold flow

1 oxidizer cold flow

2 hot fires

3 aborted tests

-Rev 2.1 fluids: panel mounted propellant feed, introduction of flow meters, panel mounted propellant valves and purge system

-Rev 2 ground electronics and software

-New concrete pad for tank stand

April 30, 2016

4 hot fires

-Rev 2.2 fluids: increased propellant line diameters, optimized flex lines on run tanks

-New blast shield

-New water cooled flame diverter and firex system

-New APCP igniters

The last test of lotus resulted in the copper chamber liner being expelled from the aluminum chamber sleeve. This was from investigations into the length of ox lead needed to get the engine to ignite compounded by inconsistent valve timing. With nitrous oxide as an oxidizer, a precisely timed oxidizer lead is needed to allow the oxidizer to decompose in the chamber due to contact with the burning igniter. With too much fuel being present at the same time, the oxidizer is unable to decompose well enough to allow combustion. The film cooling ports at the throat of the engine make proper propellant timing difficult, which led to combustion outside the chamber during most of the earlier testing. Combustion in the chamber on the first two tests may have been achieved due to the high hydraulic resistance of the original fluids system leading to lower flow rates. Less propellant entering the chamber means the igniter can more easily decompose the oxidizer, even with fuel present, preventing ignition from being smothered.

"Flamethrower" behavior from combustion occurring outside chamber

"Flamethrower" behavior from combustion occurring outside chamber

With each test, the oxidizer timing was being incrementally advanced until proper ignition could be achieved. However, due to a slow valve from trapped pressure and oxidizer bleed-in issues, the timing was incorrect going into the last test and the oxidizer was too far advanced. This led to detonation in the chamber upon fuel entry, creating over-pressurization and failure of the retaining ring holding in the copper chamber liner. This is the designed failure mode for the chamber, since all components that are ejected are comparably low energy and are deflected in a safe direction.

To address the problems that led to the chamber failure, an ox bleed valve is going to be added so the oxidizer will be starting at a consistent place in the feed line, right behind the main valve, with each test. The main oxidizer valve is going to be swapped for a vented valve to prevent pressure buildup leading to inconsistent valve opening times.

Other potential changes are to the engine itself, focusing on making the engine easier to ignite. One is increasing the L* of the engine. L* relates the volume of the chamber to the throat area, with a larger L* increasing the dwell time of the propellants in the chamber. A longer dwell time in the chamber means the propellants will have more time to react. More planning and redesign is going to occur over the summer to allow manufacture and testing to begin as early as possible in the fall semester.


To test Lotus, the existing hybrid test stand infrastructure was modified to better suit the liquid engine, and a new ground fluids system built. As mentioned in the test breakdown, there were two major revisions of the ground fluids system.

The first had many of the components and plumbing removable from the stand.  This was found to be subject to high line loss, so the line lengths were shortened where possible.  The setup was still cumbersome, so the pressurization system was built onto a panel, followed by the feed system with the addition of flowmeters in the second revision of the fluids system.

Even after this reorganization, the line losses were still higher than were acceptable, forcing the run tank pressures very high. This led to the upgrade of the feed lines again to a larger diameter. The end result is that the fluid system performs to the requirements of the engine and is fast to set up and break down in the field.

Along with the test stand and fluids system, a fire suppression system was added to the ground equipment this year. The initial setup relied on a pressurized water tank, and a pyrotechnic valve turning on the supply to the nozzles. This was then changed with an upgrade to the blast deflector. A larger tank supplies water to an electric pump, which can be controlled by the ground support electronics. A water spray cools the blast deflector during normal operation, but in the case of a fire needing to be extinguished the flow is switched on to nozzles mounted on the test stand.


With the new requirements presented by Lotus, the ground electronics and software had to be upgraded for testing as well.

The software was substantially rewritten focusing on both increasing its capabilities and reliability. The autosequence and abort handlers were changed to allow them to be more flexible and easily modified, including configuration files that could be changed with the software running, without having to reset it. This increases testing pace substantially as the ground support system does not have to be safed and shut down for the reset like it was before. The abort handler was made more reliable and predictable, and not in need of being reset after being triggered. This allows faster recovery of manual control after an automatically triggered abort.

Usability was also upgraded, including an improved UI with window configuration saving. The UI also changes to reflect abort states for easier recognition of fault scenarios and the safe recovery from them. The software is also capable of running reliably across multiple OSs.

Example screenshot of BURPG's ground control software

Example screenshot of BURPG's ground control software

On the hardware side, the Ground Operation Devices (GOD) Box underwent a significant upgrade. Originally designed for Hyperion Rev A, the front panel and wiring harness was not capable of taking advantage of Hyperion Rev B’s full capabilities.  A new front panel was made allowing all channels to be broken out, along with new features like power control for the fire suppression pump. A new wiring harness was made with shielded wire bundles, further lowering the noise floor on measurements. The design of the new front panel and harness also allows easier access and repair to the devices in the GOD Box.



While Lotus development and testing was indeed the main focus for BURPG this year, there was further work on the Starscraper airframe and avionics. Most of the airframe work from this year is still in the concept stage, and relates to modifications that need to be made to the original hybrid Starscraper airframe for the new liquid iteration. The avionics too were revamped for the requirements of the new liquid engines.

Airframe work focused on 3 main components; the aft structure, the fuel tank, and the intertank bay, which is the area between the fuel and oxidizer tanks.

The aft structure contains the joint that allows a pair of electromechanical linear actuators to gimbal the engine cluster, and is intended to function as a stand-alone assembly to allow lab testing and characterization of the actuator control loops.

There are two different fluids scenarios being investigated for use on the aft structure. One relies on more traditional flexible tubing for the fuel and oxidizer feeds, which while based off of commercial off the-self equipment and is a lower risk design, has higher line losses and lower performance.

Aft structure concept utilizing commercial flex lines

Aft structure concept utilizing commercial flex lines

The other option being investigated is a custom bellows joint for the ox feed, since this is the higher flow propellant and more susceptible to performance issues due to inefficient feed arrangements. The bellows joint would comprise of a rolled metal bellows in the center of a ring joint.

Bellows joint for oxidizer feed system

Bellows joint for oxidizer feed system

Another item of the airframe being worked on is the fuel tank. The fuel tank follows a similar design used originally in the ox tank, with modified bulkheads. The bulkhead design was changed to be lighter, more inexpensive to manufacture, and have the necessary ports for the ox pass-through, instrumentation, and pressurization inlet and tank outlet. The bulkheads on either side of the fuel tank are intended to be identical for faster manufacture, since there will be no change in tooling for each bulkhead. Final volumes and operating pressure will be determined by engine operation parameters verified through ground testing.

Concept model of fuel tank, showing internally mounted pressurization tank

Concept model of fuel tank, showing internally mounted pressurization tank

The final airframe component being designed is the intertank, which is also an important structural area of the rocket since it transfers load from the fuel tank to oxidizer tank. Housed in the intertank are critical avionics components including the inertial measurement unit and flight computer, as well as control hardware for the fuel tank pressurization. All of this equipment needs to be accessed easily. The structural members of the intertank are struts placed radially around the bay, and are removable for servicing the equipment installed in the bay. There is also a central, non-structural mount for the avionics.

For the avionics on the liquid version of Starscraper, the decision was made to create a new avionics framework that would offer increased flexibility and performance through the use of modular components networked over Ethernet. These components are:

·         Linear Actuator Controller (LAC), which drives the brushless motors used in the linear actuators that gimbal the engine cluster

·         Data Acquisition and Control boards (DQCs) several of which collect data and control valves across the rocket

·         Flight Computer and Switch (FCS) which handles the major navigation calculations and serves as the central communications node of the avionics network

·         IMU board (IMU) which interfaces the internial measurement unit to the avionics network

·         Telemetry Interface Board (TIB) which allows the telemetry system comprised of COTS components to interface with the avionics network

While a modular, Ethernet networked avionics framework is new for BURPG, the avionics components rely on mature technologies developed by BURPG over several years of avionics improvements. These boards are in various stages of development. The LAC has already undergone debugging and preliminary linear control loops have been tested. The DQC is undergoing assembly, with the other boards having completed the layout stage.

Work also began this year on a rebuilt ASTRo, used as a testbed for inertial control system implementation. The ASTRo airframe was rebuilt with minor modifications to allow more reliable recovery and more secure camera mounting. The most drastic change was to the flight computer, which was redesigned to utilize a higher performance IMU; the same one intended to be used on Starscraper. The ASTRo flight campaign will happen during the 2016-2017 school year following the completion of a first revision control loop.

New ASTRo flight computer with IMU

New ASTRo flight computer with IMU



Over the summer, BURPG will be working on several things to prepare ourselves for Lotus testing as early as possible in the fall semester. This work will largely focus on the changes to Lotus and test infrastructure as highlighted earlier. Concept development for the vehicle will continue, as will work on ASTRo's control system. The plan is to begin vehicle integration with the start of the spring 2017 semester. Next stop is space!

Beginner Rocketry Resources

Many think that rocketry is complicated (it is rocket science, after all), but in fact rocketry is quite easy to understand when broken down to its core principles; it's breaking it down that can prove daunting. One of our rocketry friends across the river at MIT, Zachary Bierstedt, recently published a great article covering some key rocketry concepts. It is a very good primer, covering everything from the physics at play through to the major components of a rocket and their function. 

For those of you curious what's going on when us rocket nerds get down to business, please check out the article linked below.

Introduction to Rocketry



Also, for some robotics resources, see the link below. Robotics is another interest of many of our members, and has a lot of overlap principles; especially where electronics are concerned: 

Robotics Education Resources




Lotus Engine Hot Fire 3

The weekend of April 2 saw the first test of Lotus with a corrected triplet injector. From the previous ignition tests 1 and 2, it was clear that a known design flaw had reduced the injector stiffness to the point where the feed system was able to couple to the engine, resulting in oscillations much lower than any acoustic chamber modes.

The new injector attempted fixed these issue with properly sized fuel ports. Unfortunately, the day’s events did not provide useful data on engine performance. However, the day was a good shakedown of a heavily upgraded fluids system and an exercise in rapid response to operational changes.

Upgraded pressurization system and oxidizer fill panel

Upgraded pressurization system and oxidizer fill panel

Testing was to start with a 2 second ignition test, followed by a 5 second burn to obtain regenerative cooling data. The first attempt resulted in a manual abort being called due to non-nominal tank pressurization. The pop valves had been moved closer to the bang-bang pressurization solenoids on the upgraded fluid system, meaning they saw higher pressure than before and were opening. The tanks run at 900 psi, but the pressurization line immediately after the solenoid valve was exceeding the valves’ rated 1000 psi. Due to the amount of GN2 vented through the pop valves, the GN2 battery pressure dropped from 2700 psi to 1500psi in under 5 seconds.

It was determined that the test could be run without the pressure-relief backup of the pop valves due to the automatic aborts programmed into the ground software, so the valves were removed. Due to the loss of GN2, a conditional abort was added to shut down the engine if GN2 pressure dropped too low during a run.

Draining the oxidizer tank during a system-safe abort

Draining the oxidizer tank during a system-safe abort

Everything was reset for a second attempt, which ended on an automatic system triggered abort. An existing conditional abort to shut down testing on oxidizer/fuel ratio was triggered early with the activation of the GN2 pressure abort flag. The oxidizer/fuel ratio abort then shut down the system, since the not-yet-ignited engine was not providing nominal data. The system went into system safe while the abort configuration file was quickly rewritten to remove the oxidizer/fuel ratio abort, being deemed unnecessary for a 2 second ignition test. With the corrected aborts in place, control of the system was regained and testing reset for a second time.

Combustion occurring outside the chamber

Combustion occurring outside the chamber

The third attempt of the day went smoothly operations-wise. However, due to a check of the main fuel valve that was run after the fuel tank had been filled, the fluid level in the fuel feed tube was different than normal. This advanced the fuel timing, which led to fuel and oxidizer entering the chamber at the same time. Proper ignition of the engine relies on the oxidizer having decomposed when exposed to the pre-burner ahead of fuel entering the chamber, which did not happen effectively in the presence of the fuel that was entering early. This caused the mixture to ignite outside the engine once mixed with atmospheric oxygen, creating a very fiery, low thrust burn. Video footage seems to suggest that the fuel/ox mixture was not igniting until after leaving the chamber.

Data Analysis and Results:

Due to the combustion occurring outside the chamber, the engine performance was obviously not valid. Due to not having choked flow, the thrust data was very low. The thrust level matches that expected by the pure momentum of propellant discharge.

All the operating pressures were non-nominal as well. The ideal are as follows: 480 psi oxidizer manifold, 660 psi fuel regen inlet, 460 psi fuel manifold, and 400 psi chamber. From further analysis of the fluid system, the line losses appear to be higher than initially expected. Taking the data seen here, the lines are being upgraded appropriately to decrease pressure drop to an acceptable level. A filter on the fuel system is also being removed, and replaced by a filter on the tank fill inlet.

Along with the upgraded fluids lines, there are operational changes planned. A fuel bleed process is being added, to make sure the fuel level in the feed lines is consistent from test to test. The software has been upgraded to make the abort system more predictable and flexible, as well as easier to modify.

The water suppression system is also being changed. Instead of a pneumatically pressurized deluge tank, a larger plastic tank and pump is going to be used. The water suppression also includes a new blast plate with water channels for plume suppression and to prevent severe heating of the plate. 

Lotus Engine Hot Fires 1 and 2

Feb. 23, 2016

Early afternoon on February 20, 2016, BURPG successfully completed two ignition tests of the Lotus engine. The data obtained was highly sought after, showing engine performance and operating conditions. This is critical for further development of the engine for flight readiness, performance, and reliability.  
The first test of the day was a final cold flow on the fuel system to verify previous changes had the desired effect.  After a quick reset, the first 2 second hot fire was attempted, which resulted in a hard start and an abort 1 second into the burn.  After inspection and some valve timing changes, a second 2 second hot fire was attempted. Ignition was much softer, though the support systems again threw an abort 1 second in. These aborts triggered as expected given the engine operating conditions, which was a good verification of the ground support system capabilities. 

Lotus on the test stand ahead of the test

Lotus on the test stand ahead of the test

Fuel Cold Flow 

The fuel cold flow had the objective of checking expected pressure drops across the system as well as getting valve timing for the ignition sequence. 

Pressure drops were within expected ranges, and the time between the valve opening and fuel entering the chamber was calculated. This data is used for the valve timings for the hot fire.

Hot fires 

Lotus during hot fire testing

Lotus during hot fire testing

Shortly after noon, preparations were complete for the first hot fire.  Fill procedures were swift and smooth from the control station.

With ignition, there was a loud bang indicative of a hard start. An automatic oxidizer to fuel ratio abort also triggered upon being armed about 750 ms into the burn. Inspecting the engine after the test, it was clear that there was no damage. Pressure data from the ignition was used to alter the ignition sequence to introduce the propellants in the chamber at almost the same time in the next test with a slight oxidizer lead. Pressure data also verified that the ratio abort properly triggered due to the ratios being out of bounds.

The second ignition was much more gentle, with a much more desirable startup transient. Despite the better ignition, the ratio abort triggered again due to out of bounds data. This was not unexpected from the previous test’s data.


The following data comes from the second hot fire of the day.

The engine achieved 66% of the target thrust, which oscillated during the burn. The cause of the initial spike in thrust and the subsequent variance can be seen when looking at the chamber pressure. The actual oxidizer to fuel ratio the engine operated at is still in question due to these fluctuations disrupting the calculated mass flow data.  To correct this, new flow meters will be added to our system in place of using tank mass to calculate mass flow. The flow meters will offer better data independent of vibrations and pressure oscillations from the engine.

The propellants entered at the correct time relative to each other, but ignition still happened about 100 milliseconds too soon. The igniter timing will be adjusted to correct this. The previously mentioned variation in thrust and chamber pressure begins after the first 100 or so milliseconds of operation. This is combustion instability, meaning waves of combustion are travelling through the chamber. This pressure variation is what causes the buzzing sound in the video as the engine fires.
The cause of the instability appears to be in the injector. A key parameter of injector design is “stiffness,” which is the ratio of the pressure drop across the injector and the chamber pressure.  The stiffness on the fuel injectors was supposed to be 15%, but in this test ended up being more along the lines of 5% due to initial errors in fuel density calculations.
A new injector is already being made to increase stiffness, which should also help prevent instability as well.

Finally, shutdown of the engine was successful, though slow. This leads to the fiery shutdown seen in the video. This behavior was expected due to the fuel left in the regenerative cooling channels.

Above is data on the regenerative cooling system.  The engine was far from steady-state at only 1 second into the burn, but it’s very gratifying to see decent temperature data!  The upcoming longer burns will provide better data on regen performance and steady state engine operation.

Specific impulse is a measure of the efficiency of an engine, measured in seconds.  The number provides the ratio between the thrust obtained and the weight of fuel per second needed to maintain that thrust.  The engine did not reach the calculated 235 seconds. This is likely due to the aforementioned combustion instability as well as a skewed oxidizer to fuel ratio from the outgoing injector design. The specific impulse is expected to increase significantly with the corrected injector, which will allow the engine to operate much closer to the desired ratio.


To summarize, these tests provided very valuable data on the engine’s performance and behavior and answered some outstanding questions on fluids systems performance that could only be answered by igniting the engine. From the data, we were able to identify areas to improve on in the next few weeks. These largely focus on making a new injector to correct the oxidizer to fuel ratio and increase stiffness, which will in turn lead to more stable combustion, higher thrust, and much better specific impulse.  Flow meters will also be added to the ground support system to provide better mass flow data, leading to more exact impulse and ratio  calculations.

Stay tuned as we improve Lotus for the next round of testing!

Lotus Liquid Engine Cold Flow #1

Dec. 5 2015

BURPG has completed the first successful cold flow of the Lotus liquid engine, obtaining valuable data on the new ground fluids system performance, valve timings, and preliminary engine parameters. It was a significant first step in the development of the Lotus engine and further progress on BURPG’s way to space.

The testing included 4 oxidizer flows and 3 fuel flows. These varied in length, with the longer flows providing data on engine behavior, and the shorter flows providing a larger data set on valve timings. Determining these timings are critical for the engine ignition autosequences. Following is a brief analysis of our results and how these will be applied to future testing and revision.

See the video of the test here:  https://youtu.be/oEC2i0adbSg


Pressure drops are important engine parameters that can be found during cold flow testing. The lotus pressure drop data provides some interesting insight into the engine behavior during the cold flow.

Below is a graph of the fuel system pressure drops from a long duration fuel flow. The regen dP is the pressure drop across the regenative cooling channels, and is determined by the difference between the regen manifold pressure and the fuel injector manifold pressure. Other pressure drops are determined by measuring the pressure relative to atmosphere. Before the main fuel valve (MFV) opens, there is pressure present from the purge system. The purge check valve opening is also responsible for the small spike after the main fuel valve closes.

For the regen pressure drop, we were expecting 200psi, and the fuel injector 60psi. The low values shown in the graph are caused by higher than expected losses on the ground fuel system, and can be corrected for by increasing the test tank pressures. The mass flow rates that were calculated also came in at about 75% of what was expected, in agreeance with the lower than expected pressure drops.

The graph below details oxidizer injector pressure drop. As noted with the fuel, the purge pressure is seen again, this time as the pressure in the regen system and fuel injector. Purge pressure in the fuel system increases during oxidizer flow as the check valve for oxidizer purge shuts, diverting more flow to the fuel purge.

For the oxidizer injector, the expected pressure drop of 80 psi is actually lower than measured. The oxidizer mass flow rate was also calculated to be about 37% of what was expected. While this seems counterintuitive to have lower than expected flow rates and higher pressure, this can be explained. During cold flow conditions, the vapor pressure of nitrous oxide is significantly higher than the chamber pressure, causing the liquid nitrous oxide to cavitate and turn to gas in the injector ports, restricting the flow area and leading to the behavior seen in the data. Since this data does not verify the pressure drops of the oxidizer injector, other methods can be used. The CdA of the injector can be verified through water flow testing, from which the approximate pressure drop can be backed out.

Visible in both graphs is the effect of the pressure swings from the pressurization system, creating fluctuations in regen pressure and oxidizer injector pressure. More discussion on the ground support equipment performance follows.



Since we were using an entirely new ground support equipment (GSE) system for the liquid engine, the cold flow was as much a test of the GSE as the engine itself. The GSE system performed well for a first test, but observations and data from the test highlight some areas to focus on to make the system even more effective.

One area is the pressurization system. The press system is intended to keep the test tanks within a 15psi margin around the desired set point as the tanks drain by using a bang-bang solenoid valve controller and 3000psi nitrogen. The pressurization controller also utilizes the tank vent valve in the case that there is significant overshoot on the tank pressure to limit pressure fluctuations making it to the engine.

From the pressurization data, it is clear that the pressure swing was very significant. This is from two factors. One was the software buffer for the data was not being flushed frequently enough, making the response time of the controller in excess of 150ms. This slow response was compounded by a second factor, the extremely high pressurization flow rates through the press system. As evident in the video of the test, the pressurization swings were large enough to consistently trigger the vent valves. The pressure swings are evident in the graph of the oxidizer test tank pressure below. The dotted lines represent the set points for the controller, and the solid line the goal pressure.


These two issues are very straightforward to fix. By flushing data buffers more frequently, the response time of the system can be brought down significantly. Response time can be further improved by hard coding the controller onto the DAQ and control board, Hyperion. Combining this with orifices on the nitrogen feed to limit flow, the bang-bang controller should be tunable to the point that the pressure swing is within our set points and the vent would never need to open.

A second fix that is needed is to the vent valves themselves. On the oxidizer tank, the valve points down the side of the tank. This orientation allowed the thrust from the valve to actually push up on the oxidizer tank load cell, creating fluctuations in weight data with the valve opening. On the fuel tank, a horizontally mounted vent valve was able to swing the tank, as evident in the video. To remedy this, both valves will have a T fitting attached to the output, so that there will be no net thrust from the vent valve. The tanks will also be constrained in the stand from horizontal movement. An example of the effects the vent valves had on mass measurement of the fuel tank is given to the right.


Because of the fluctuations in the tank mass during the test, the mass flow rate had to be backed out from the starting and ending mass versus the time the test ran. Since this is not ideal, future testing will focus on obtaining more accurate mass flow data.



For proper ignition of the engine, oxidizer and fuel have to enter the combustion chamber at precisely the right time. This is made difficult by the fact that the fuel has to make it through the regenative cooling channels before reaching the injector. To determine the time from valve open command to arrival of each propellant at the injector and valve close command to the end of propellant flow, the valve command was overlaid on the pressure data from multiple tests. Below, an example from an oxidizer flow test is shown. The difference between the valve command (orange) and the pressure (blue) starting to change was taken as the valve timing. For ox and fuel, start up timing was remarkably consistent, with all tests being within 3ms and 12ms of each other, respectively. Shut down timing was more variable, varying as much as 27ms and 40ms, respectively.

valve timing.PNG



For future testing, the GSE system will implement fixes for the issues described above, providing better mass flow data and less fluctuation in feed pressure. The next test will be a cold flow to verify our GSE corrections. The pressure drop across the oxidizer injector will also be roughly verified by water flow testing. Following the completion of these two steps, hot fire testing will begin. Hot fire testing will start with several short duration tests, focusing on measuring start-up transients and ensuring that the engine can be ignited safely and consistently. Testing will then move on to longer tests focusing on measuring steady state performance as well as varying oxidizer to fuel ratios.


BURPG Presents at Embedded Systems Conference in Silicon Valley

Two members of BURPG, Dean De Carli and Matthew Owney, presented at the 2015 Embedded Systems Conference this past week. The presentation, titled Tire Rubber, Laughing Gas, and Silicon: A Rocket Story, touched on the struggles, triumphs, and technical details of designing a complete avionics and support system for the worlds most advanced amateur rocket from the ground up. To read more about the presentation, please see the EE Times article


Mk. V Hot Fire #1 Results

On May 2nd, 2015, BURPG attempted the first hot fire and full systems test of the Mk. V hybrid rocket engine, the most powerful amateur rocket engine ever built. The hot fire was intended to test the throttling profile to be used during flight, as well as characterize the TVC response. It would also verify the durability of the combustion chamber and nozzle. Following the test, the results would be used to optimize throttling for maximum altitude, and adjust gains in the TVC control algorithms to obtain stable flight.

The hot fire test did not go as planned. A software crash, and ensuing abort, led to valves not operating as intended, causing the combustion chamber to rupture. A full breakdown of the events during the test follows.

During fueling, the telemetry and control stream between the control station and the ground support electronics was lost. This triggered a system-safe abort, which functioned as intended. All fueling valves closed, and vent valves opened to start releasing pressure, in the event that contact could not be reestablished. Connection was reestablished and lost propellant replaced. At this point, a TVC sweep test was done, to verify valve control before the hot fire began. During this, one of the TVC servos stopped working, leaving the TVC valve open. This was seen as a non-issue, since its effects would not interfere with the rest of the TVC sweep test during combustion, and the nitrous oxide ice forming on the nozzle would be burned off during the ignition sequence by the long preheater burn.

The autosequence was initiated, and the preheater burned as intended during the ignition process. The pyrotechnically actuated valve then opened, allowing the oxidizer to enter the combustion chamber. However, the engine stayed at idle. This was due to the throttle valve not opening for the ignition process, but rather remaining closed. It is not designed to seal fully, so oxidizer did leak into the combustion chamber. The cause of the throttle valve not responding still remains to be determined, though it was not something that came up in normal testing. Most likely explanations stem from some effect of the software crash.

After about two minutes of the engine idling, a burn through was noticed in the post chamber, directly above the nozzle. A combustion abort was called, which would shut the throttle valve. The throttle valve instead opened, flooding the combustion chamber with oxidizer. The flame near the base of the combustion chamber ignited the engine from the bottom up, over-pressurizing the chamber and causing it to rupture.

The safety shield on the trailer kept all propellant bottles safe. The fire department was on-call, so the burning combustion chamber was extinguished within minutes. All that was lost in the test that was not disposable were the TVC valves, servos and bottom combustion chamber bulkhead. The trailer was undamaged, except for scorching on the safety shield. Due to the way BURPG operates the rocket fueling and tests, no one was in danger at any point. All operations are handled from a safe zone more than 700 feet away from the rocket from the moment the propellant tanks are opened.

Anyone familiar with rocket engine development is aware that engine failure is always an option. The only way a test can be a failure, however, is if nothing is gained from it. While the first hot fire of the Mk. V may not have been successful in terms of testing the engine performance, it is important to still reflect on what was gained. Our members now have invaluable experience in designing, assembling, and testing of highly complex rocket engines, on par with what is seen in professional aerospace. We also identified problems and their solutions. New networkable digital servos, with feedback, are now going to be used on the rocket. With the addition of feedback, servo position can always be verified, allowing better diagnosis of the system and closing the loop on the controls. A weak point in the software is also being massaged out to provide better reliability.

Moving forward, the Mk. V will be rebuilt and retested. A second combustion chamber is nearly complete. TVC valves, servo brackets, and combustion chamber bulk plate are all being redesign and remachined over the summer. Upon returning to school, fuel grains need to be cast and a new nozzle made. We would then make a second attempt at a hot fire, and hopefully be ready for a flight following that.

All of us here at BURPG are incredibly excited to continue the Mk. V campaign, and all look forward to the day this thing flies. We thank the Boston University College of Engineering for their continued support and encouragement, as well as our many sponsors and backers. It has been a lot of work to get to this point, but incredibly rewarding for all involved. We are excited for what the future holds for this project.

Rockets are high risk, high reward. We all understand that; it’s also what makes all of this so much fun.