Claybaugh 6-inch Rocket, Final Report

by Bill Claybaugh, RRS.ORG


EDITOR’S NOTE: This is a continuation of the reporting from the 10-16-2021 flight of the 6-inch rocket design, built and flown by RRS member, Bill Claybaugh.


This project is part of an effort to develop a two-stage sounding rocket capable of sending about 5 kg of usable payload to about 200 Km altitude.  This vehicle is intended to act as the upper stage of that two-stage rocket; it was—based on a systems analysis–sized for an eight second burn-time and about 1300 lbf thrust.

OVERVIEW

As flown the vehicle was 101.25” from nose tip to the fin trailing edges.  The Payload section was 40.125” in length and 6.170” in diameter; the booster was 61.125” in length and 6.00” in diameter.

Computer simulated rendering of the rocket

The vehicle had an aluminum nose tip, a filament wound fiberglass nose with a 5.5:1 Von Karmen profile, a filament wound cylindrical payload section, and an aluminum motor / airframe with aluminum fins.

VEHICLE DESCRIPTION

Inspection of the Forward Bulkhead showed it to be in good condition with no evidence of any gas leaks above the two O-rings.  The bottom of the Bulkhead showed some damage to the fiberglass heat shield from the ground impact of the rocket but showed plenty of

Pre-flight estimated motor performance was 1350 lbsf. of thrust with a burn-time of 8.35 seconds.  Burnout velocity was estimated at Mach 3.1 at 14,400 feet with a peak altitude estimated at about 71,000 feet. Total flight time was expected to be 143 seconds.  The booster had a streamer attached at the forward end to try and cancel horizontal velocity upon deployment at peak, thus limiting the range of the booster.

The payload also used a streamer for recovery, it was planned to separate from the booster near peak altitude using a pneumatic separation system that operated four pins which rigidly attached the payload to the rocket until pressure was released.

Final vehicle mass properties are shown below:

 Item DistanceWeightMomentC.G. 
   from  from 
   Nose Tip  Nose Tip 
        
 Nose Cone Assy. 24.005.55133.20 Measured
        
 Ballast 29.900.000.00 Estimated
        
 Instruments Assembly 33.008.97296.01 Measured
        
 Bulkhead & Sep. Sys.: 37.502.80105.00 Measured
 Bolts 38.630.031.28 Measured
        
 O/A Payload Length: 40.13    
        
    17.35535.4930.858 
        
 Bulkhead Retainer 43.041.1047.34 Measured
 Bolts & Nuts 43.410.3414.59 Measured
        
 Bulkhead Assy. 40.192.3092.43 Measured
        
 Tube 69.4313.05906.00 Measured
        
 Outer Liner 68.314.10280.08 Measured
        
 Fin Can 95.102.75261.52 Measured
        
 Nozzle 97.497.83763.31 Measured
 Bolts 97.190.2221.38 Measured
        
 Fins 96.495.55535.53 Measured
 Bolts 96.190.1816.93 Measured
        
 O/A Stage Length: 101.25    
        
    54.773474.5963.445 
        
 Propellant 68.3154.203702.54 Measured
        
    108.977177.1365.866 
        

FORWARD BULKHEAD

The forward bulkhead assembly consisted of the forward bulkhead with O-rings, fiberglass spacers, and a bulkhead retainer that incorporated the bottom portion of the separation system (four holes for the attachment pins and a 45-degree bevel to allow the payload to fall off the booster once the four pneumatically operated pins retracted).

Computer rendering of the forward bulkhead
Bulkhead retainer with separation system fittings

FINS

The Fins were attached to the motor tube via an internal “fin can” that served to provide the “meat” to allow four countersunk fasteners to hold each fin rigidly to the motor tube.  The internal Fin Can had a single O-ring at the top to seal between the phenolic propellant liner and the fin can as wall as two O-rings to seal between the fin can and the motor wall.

Computer rendering of the internal fin can

Note that the fins shown are the flight fins, post flight; with the exception of minor gouging the fins appear to be fully reusable.

Photo of fins post-flight
Computer rendering of one fin

NOZZLE

The nozzle consisted of an aluminum outer shell, a graphite insert, and a stainless steel nozzle extension with a plasma sprayed Zirconia overcoat on the inside diameter.

Photo of the nozzle
Photo of the nozzle from the other side

PROPELLANT LINER

The liner protecting the motor tube from the combustion gas was a phenolic tube with a 5.50 inch inside diameter.  The tube was originally slightly oversize for the motor tube’s 5.75” nominal inside diameter and was sanded as necessary to make it a tight slip fit into the motor tube.  It was then cut to a 48” overall length and fitted to the motor tube using a high temperature grease (550 degrees F).

Post-flight analysis shows that the liner had about 0.090” – 0.092” of the original 0.125” wall remaining in those areas exposed to hot gas throughout the burn; note that heating of the phenolic leads to expansion of the thickness of the liner, nonetheless, there was no evidence of hot gas having reached the motor tube wall.

PROPELLANT

The grain was cast in place using a dissolvable (polystyrene) mandrel that provided for four fins at the base of the motor and a simple cylindrical core at the upper end.  This grain design provided an approximately neutral thrust curve as the finocyl section regressed in burn area at a rate that very closely matched the progression of the cylindrical section of the grain.

Grain cross sections
Thrust and chamber pressure curves

The finocyl section at the base of the grain was 14.75” in length, the cylindrical section 31.25” in length for an overall 46” propellant grain length.

Because the grain design tools used for this project worked only in two dimensions, the 2.66 square inches of exposed grain surface at the top of the finocyl fins was not modeled in the simulation.  This represents 0.80% of the initial grain burn area and, accordingly, the actual performance was expected to be slightly regressive.

All grain design simulations were based on the 0.056 lbsm / cubic inch propellant density of the various static test motors; in the event, this grain came in at 0.059 lbsm / cubic inch due to changes in both the propellant mix and processing.  The effects of that higher density on flight performance will be addressed in the Analysis section.

AERODYNAMIC MODEL

Most dynamical simulations for this flight were conducted using RASAero II. The aerodynamic model estimated by that tool is shown below:

Aerodynamic model plot

Likewise, RASAero II provided estimates of Stability Margin over the flight profile:

Stability margin plot

A splash analysis was very graciously conducted by Chuck Rogers.  That analysis concluded that the initial launch conditions that minimized risk to the uninvolved public were a launch azimuth of 244 degrees and a launch tower angle of 87 degrees (that is, three degrees below vertical in a southwesterly direction).

PAYLOAD

The payload consisted of three subsystems: a pneumatic payload separation system, a main flight computer with integrated transmitter, and, a backup flight computer with onboard recording of flight engineering data.

PNEUMATIC SEPARATION SYSTEM

The separation system relied on four pins that rigidly locked the payload to the vehicle. The system was actuated by command from the main or backup flight computers, which command fired a nitrocellulose-based initiator that in turn drove a plunger through a burst disk.  Venting of the system allowed spring force on the four locking pins to draw them inward, thus allowing the payload to fall away from the booster.

The Separation System was o-ring sealed at all connections to assure it remained leak free under flight conditions. Initial testing showed the system could hold pressure (125 psia air) for 100 hours.  Pre-flight testing included a 50-hour leak down test followed by one minute on a shake table.  The unit was leak free and actuated on command after this final test.

The main flight data recorder and transmitter was a Multitronix Kate 2 System; backup flight data recording was provided by an Altus Metrum EasyMega.

Photo of the locking pin system
Photo of the pneumatic separation system

MAIN FLIGHT COMPUTER

The main flight computer was a Kate 2 Data Recorder and Transmitter from Multitronics, Inc.  This system used a 915 MHz ISM uplink and downlink with on-the-fly adjustable power output from 100 mw to 1 watt, it used Spread Spectrum Frequency Hopping and FSK Modulation with a 128-bit AES encryption.

The system fixes its GPS position every 200 msec and features unlimited GPS altitude reporting; the velocity lockout is at 1700 ft/sec.  A 50 g Axial Acceleremeter and 10 g pitch and yaw accelerometers record every 10 msec and report via telemetry every 100 msec.  A separate pyro board initiates payload separation and peak.

The transmitter link budget indicates a worse case net 26.5 dB at the receiver for this flight.

Link budget details for the flight computer transmitter

BACKUP FLIGHT COMPUTER

The backup flight computer was an Altus Metrum EasyMega with three axis data recording (acceleration and rates) and a barometric altitude estimator.  Separate batteries and switches powered the independent pyro initiation which was programed for one second after the accelerometer measured peak altitude.

FLIGHT SIMULATION MODEL

Simulation using RASAero II showed an estimated peak altitude of about 71,000 feet, a worse case total flight time of about 144 seconds (assuming no separation at peak), and a maximum worst-case range of about 75,000 feet.

Simulated trajectory plot

Baseline flight simulation (from RASAero II):

Baseline flight simulation

Launch Angle Vs. Range (from RASAero II):

Simulated launch angle vs range

Maximum Range Estimation (from RASAero II):

FLIGHT TEST RESULTS

Based on video analysis, ignition require 0.067 seconds from the rupturing of the burst diaphragm (a standard 1.5” rubber stopper previously tested to pass the nozzle at 40-50 psia) to first motion. From first motion, it required 0.35 seconds to clear the 24-foot tower at about 25 feet altitude and about 165 ft/sec.

Frame-by-Frame Video Analysis

(Red Indicates Clearing the Tower)

Estimated
CumulativeEstimated
FrameEstimateEstimateEstimatedEstimatedAverageInterval
NumberBurnFlightVerticalVerticalVerticalVertical
TimeTimeMotionVelocityAccelerationAcceleration
(ft.)(ft./sec.)(g’s)(g’s)
10.0170.0000.00
20.0330.0000.00
30.0500.0000.00
40.0670.0170.5060.00110.80110.80
50.0830.0331.0060.0054.90-1.00
60.1000.0502.0080.0048.6936.27
70.1170.0672.5075.0033.94-10.32
80.1330.0832.7566.0023.60-17.77
90.1500.1003.0060.0017.63-12.18
100.1670.1174.0068.5717.2514.97
110.1830.1335.0075.0016.4710.98
120.2000.1505.0066.6712.80-16.53
130.2170.1678.0096.0016.8953.66
140.2330.1838.5092.7314.71-7.10
150.2500.20010.00100.0014.5312.55
160.2670.21712.00110.7714.8819.07
170.2830.23314.00120.0014.9716.20
180.3000.25017.50140.0016.3936.27
190.3170.26719.50146.2516.0310.65
200.3330.28322.00155.2916.0215.85
210.3500.30025.00166.6716.2520.19
220.3670.31728.00176.8416.3417.96
230.3830.33331.00186.0016.3316.06
240.4000.35033.00188.5715.733.79
250.4170.36736.00196.3615.6313.52
260.4330.38340.00208.7015.9121.98
270.4500.40042.00210.0015.301.43
280.4670.41746.50223.2015.6423.60
290.4830.43354.00249.2316.8647.50
300.5000.45057.00253.3316.486.64
310.5170.46761.00261.4316.4014.08
320.5330.48364.50266.9016.159.19
330.5500.50069.00276.0016.1415.96
340.5670.51773.00282.5815.9911.26
350.5830.53376.50286.8815.707.00
360.6000.55085.00309.0916.4540.40
370.6170.56790.00317.6516.4114.94
380.6330.58394.00322.2916.167.64
390.6500.60097.50325.0015.824.06
400.6670.617102.50332.4315.7412.85
410.6830.633110.50348.9516.1129.77
420.7000.650115.00353.8515.918.13
430.7170.667119.50358.5015.707.67
440.7330.683123.50361.4615.434.52
450.7500.700128.00365.7115.236.92
460.7670.717132.50369.7715.026.55
470.7830.733141.00384.5515.2926.54
480.8000.750145.00386.6715.012.95
490.8170.767150.00391.3014.857.64
500.8330.783158.00403.4014.9921.55

Just after 0.50 seconds the vehicle began an unplanned turn to the Northeast.  This turn continued for 0.25 seconds before the vehicle resumed stable flight on the new azimuth and with a flight path angle of about 75 degrees.  After 0.80 seconds but before 1.0 seconds, the telemetry failed.  The cause of this failure is not yet established but appears to the manufacturer to have been a power outage; however, the battery was still connected to the main computer after recovery and the battery tested at an optimal 3.87 volts.

At about 1.0 seconds, the payload separation system appears to have been actuated by the backup flight computer; that computer is currently at the manufacture for data extraction to try and determine why it fired the initiators.

Based on video analysis, the vehicle appears to have coned twice following separation of the payload.  This coning could have been associated with the payload separation or with the deployment of the rocket’s streamer.  In either case, the vehicle resumed stable flight (as designed) without a nose cone. The payload assembly was located about 120 feet from the launch tower on the northeasterly azimuth.  The backup flight computer was still actively reporting (via “beeps”) it’s status but the main computer was not so doing.

Launch plus 0.50 seconds
Launch plus 0.75 seconds

The booster was located north and a little east of the launch site at a range of 14,300 feet.  Based on that range and the estimated motor performance a trajectory reconstruction suggests a maximum altitude of 21,200 feet, a burnout velocity of 1550 ft/sec and a terminal velocity of about 820 ft/sec with a total flight time of about 74.5 seconds.

The booster impact left an about 2.0-inch-deep depression in the hardpan before the hardware apparently fell on its side. Given an estimated terminal velocity of 820 ft/sec, this implies and average of 410 ft/sec to stop and thus that the vehicle came to rest in about 0.000407 seconds.  This in turn indicates an average deceleration of about 31,200 g’s on impact.

ANALYSIS – THE TURN TO THE NORTHEAST

The Turn to the Northeast

All testable reasons for the turn to the Northeast after 0.50 seconds have been ruled out: there was no hot gas leak nor any apparent disturbance to the thrust vector. The wind was from the Northwest and less than 5 mph, if it had caused the turn we would have expected the vehicle to turn toward the Northwest, not the Northeast. The temporary “hanging” of a part of the bellybands appears ruled out by the absence of any gap between the fins and the motor tube as well as by the absence of any damage to the fin leading edges.  Further, the bellybands all landed within fifty feet of the launch tower; given an estimated velocity of about 165 ft/sec at the top of the tower, this implies that each bellyband followed a nearly vertical trajectory following clearing the tower.

The remain hypothesis for the cause of this turn is that the vehicle ran into a “dust devil” that was not visible because it had not reached the ground.  Examination of the video using polarized glasses showed no evidence for such an event, but that is not conclusive as the sun angles may have been inappropriate for this technique.

ANALYSIS – TELEMETRY FAILURE

Telemetry failed after 0.80 seconds but before 1.0 seconds based on analysis by the manufacturer of data recorded by the receiver (data packets are sent every 0.2 seconds, one was received at about 0.80 seconds and no subsequent packets were recorded).  The cause of this failure is unclear: the manufacturer has initially concluded it was a power failure, however, the battery showed 3.8 volts at recovery and was still connected to the computer / transmitter; thus, a power failure would have to have been internal to the hardware. This failure might be associated with separation of the payload from the rocket which occurred around this time.  Transmitted data show that the main computer did not initiate the separation and had continuity to the initiator throughout the period during which data was transmitted.

ANALYSIS – PREMATURE PAYLOAD SEPARATION

The payload was recovered about 120 feet from the launcher on a Northeasterly heading.  Based on the location a trajectory reconstruction suggests separation may have occurred around 1.0 seconds into the flight at about 400 feet altitude.

Given the data indicating that the main computer did not command separation while it was operating and the observation, following recovery, that both initiators had been fired (firing of either initiator ignites the other), it appears that the backup computer may have initiated the separation.  That computer is currently at the manufacturer for repairs after which we hope to extract whatever data it may have recorded, including continuity data with respect to the initiator to which it was wired.

SUBSEQUENT FLIGHT

Following payload separation, the vehicle appears to have coned twice and then resumed stable flight on the new heading.  Upon recovery, the vehicle did not have its streamer attached and we assume it was lost to aerodynamic forces during the separation of the payload and subsequent coning; however, that streamer has not been recovered and so we cannot confirm when it came off the vehicle.

Per the trajectory estimate, it appears that even with a blunt front end, the vehicle may have reached around Mach 1.35 (1550 ft. / sec.) but that estimate is unconfirmed.

Note that the video measured velocity and acceleration up the launch tower was noticeably higher than the pre-flight estimate: pre-flight, velocity at the top of the tower was estimated at about 145 ft / sec while the measured velocity just after clearing the tower was about 165 ft / sec.  This difference may be due to the higher density of the propellant as compared to the pre-flight model; assuming that the ballistic characteristics of the propellant remained the same (very unlikely) modeling of the pre-flight propellant assumptions but using the higher density indicates it would produce about 5% higher thrust at about 8% higher chamber pressure due to the higher mass flow compared to the pre-flight modeled propellant.

Modelling of the vehicle performance using the actual range and these different propellant performance assumptions does not significantly change the estimated peak altitude or velocity: the somewhat greater energy of the flight propellant is spent on increased drag as velocity approaches Mach 1.35.

SUMMARY AND FUTURE WORK

The rocket motor appears to have performed as designed, albeit in off-design flight conditions.  In the absence of any explanation for the unplanned turn to the Northeast, no changes to the motor design are planned for the next flight vehicle other than the hard anodizing of the fins to help them survive future flights to still higher velocities.

The payload assembly appears to have been commanded off the rocket motor at about one second into the flight; the reason for this remains unclear at this writing. For future flights the internal payload structure will be made still more robust to prevent the internal structural failures that did occur upon impact of the payload; some of those structures will be rebuilt in stainless steel to help move the Cg forward (this was not an issue for this flight, but will be for eventual Mach 6 burnout velocities).

Further work is required on the base of the launch tower to significantly reduce the labor required to assemble and erect the tower.

The bellybands will be modified for greater strength and spring back by moving to 1095 spring steel instead of the 2024T-3 used for this flight; in addition, the guides will be lightened both to aid travel up the rail and to mitigate against any impact damage that might occur if they contact the vehicle during separation.

MTA Launch Event, 2021-10-16

by Bill Claybaugh and Dave Nordling, RRS


This firing report will be the first in a series of three articles posted on RRS.ORG. This report will cover the launch event and preparations over many days made by RRS member, Bill Claybaugh. As the attending pyrotechnic operator for this firing event, I have summarized this work for the benefit of our readers with the permission and oversight of Bill.

Bill Claybaugh has been planning to build, load and launch a large 6-inch solid motor for many months and the first attempt had finally come to pass at the RRS Mojave Test Area (MTA) over the span of almost a week starting Tuesday, October 12 and culminating in a launch on Saturday, October 16, 2021. He had studied this project very carefully and built a great many new parts and tools from his home in Colorado. The scope of this project is quite extensive and the larger goal was to enable larger solid motor building by other members of the RRS at the MTA. The 6-inch motor was just the first in what will hopefully be a growing series of similar and larger scale solid motors.

Bill Claybaugh’s description of his six-inch rocket from his Flight Readiness Review presentation.

The predicted performance of this 6-inch single grain motor was 1350 lbf of thrust for a duration of 8.35 seconds which was expected to exceed 70,000 feet; well above the RRS MTA’s standard 50,000 foot altitude waiver. This “P” sized solid motor in this vehicle required an FAA Certificate of Authorization (COA) for this flight on the prescribed dates during daylight hours. The submission of Monte Carlo simulations of the trajectory (splash analysis) were graciously performed by Chuck Rogers (author of the RASAero II software) and a necessary part of the process to verify no significant concerns for impacting nearby populated areas or structures. Also, the FAA Class 3 rocket waiver that was granted would require the launch team to contact the relevant air traffic control 15 minutes in advance of the intended launch for final permission to proceed. A separate article discussing this subject in more detail will be coming soon.

The rocket had two streamers for a recovery system which were intended to be sufficient for easier spotting of the rocket in descent rather than provide a soft landing.

Many members of the society participated in this project over the several days needed to prepare and conduct the mixing, pouring and casting process. RRS members Dave Crisalli and George Garboden lended their time and expertise in solid motor building which led to a stellar finished product on Thursday. Several of Bill’s family and friends attended and supported the preparations for launch.

Bill Claybaugh’s four-finned rocket with an end view of the four-fin 6-inch single-grain motor loaded and ready for the nozzle installation. RRS president, Osvaldo Tarditti, talks with Bill on the morning before launch.
The forward and aft views of the nozzle assembly of the Claybaugh six-inch rocket.
Bill Claybaugh holds his payload system without the fiberglass long-ogive nosecone cover.
Pictures of the different parts of the pneumatic separation system and payload.
Ed Wranoski finishes the mating of the payload on top of the single stage solid motor checking the alignment before preparing to move the rocket to the launch pad.

Given the size of the 6-inch rocket, Bill designed and built a T-slot type of launch rail with a 24-foot length on an aluminum truss structure. The system was designed to be deployed in a green-field site and easily assembled by a small team of people. There were some challenges in getting the design to work but through the combined efforts of those at the site during the afternoon and early evening on Friday, the erecting and loading process was safely completed. Susan and Ed Wranoski both had a lot of great suggestions about getting the right placement of the come-alongs to bring the launcher up to a sufficient angle to secure it by the chains and strap anchors around the pad.

The new launch rail system will be the subject of a separate article coming later on RRS.ORG. Design improvements and substantial changes are being planned such that the next launch event will have an easier time in raising and lowering this important asset for the launching of larger rockets from the MTA.

Testing of the erecting process took place into the early evening by headlights. These operations provided valuable information making launch preparations the following morning far simpler.
Bill Claybaugh, Mike Pohlmiller and Ed Wranoski secured the 6-inch rocket by two bellybands in flyaway railguide system.

During the first launch operations of the rocket, the wireless telemetry wasn’t receiving signals. After restarting the computer and replacing the nosecone, the pyrotechnic charges in the recovery system accidentally fired due to a short. The payload system was removed, inspected and replacement pyrotechnic charges installed. After protecting the terminals from a similar short during final installation of the payload and nosecone, the telemetry system was working and the launch could proceed.

The nosecone being replaced after a quick test of the payload system.
Bill’s 6-inch rocket on the rails and secured for launch.

The launch event coincided with the launch operations of our neighbors’ (FAR). We were in constant communication to assure everyone was under cover at the proper times. The weiather was nearly ideal with very low winds the whole day. After road and air checks were completed, we prepared for launch.

Bill Claybaugh prepares for firing with RRS president, Osvaldo Tarditti, amd others ready to film and photograph the launch.
Still captured from the launch footage showing the rocket clearing the tower.
Last still picture of Bill’s 6-inch rocket before going out of view of the camera.

The initial launch was swift and powerful as the motor ignited and came to full thrust leaving the launch rail. The rocket canted to the northeast opposite the intended direction of the launch rail and the vehicle appeared to corkscrew as the motor burned to its full duration before going out of sight. The recovery system appears to have fired early as one of the streamers and the entire payload module fell back to the northern side of the MTA. The spent rocket motor casing has not yet been recovered. Bill was able to bring back the payload segment for inspection at the MTA while others continued the search for the rocket.

Bill disassembles the recovered payload system after its short descent back to the ground.
Both pyrotechnic separation charges had fired.
The antenna snapped off and was not found.
Recovered flyaway railguides showed signs of recontact from the tail fins from the sharp tears and rips seen. This is a common occurrence with flyaway railguides and they can be refurbished for the next flight.

Based on review of video footage, it appears the sudden turn uprange occurred at around 100 feet and took less than 1/4 second.  The current thinking is that the separation system depressurized, producing the side-thrust that caused the sharp turn after leaving the rail. It is assumed the telemetry loss of signal (LOS) was a result of the antenna snapping off during this sudden turn. LOS occurred at 119 feet and 425 ft/sec. About 0.25 seconds later, the payload can be seen starting to fall away from the rocket which can only occur if the system is depressurized. The payload was recovered about 300 feet from the launch tower and on the ‘new’ azimuth.

After the initiators fire–and both were fired–it would be expected that applying pressure to the quick-disconnect (QD) fitting would:

(1.) NOT result in the four retention pins extending, and,

(2.) would cause venting through the diffusers. 

That is, the burst disk is supposed to be punctured due to the piston driving the hammer through it when the initiators fired and any gas generated in the system is vented past the burst disk and through the diffusers.

The recovered flight hardware instead extended all four pins, did not vent through the diffuser, and did vent through the outlet reserved for the hot initiator gases.  This means that the burst disk was not opened and pressurizing gas was somehow leaking into the hot gas circuit.  The image below of the burst disk shows its condition as found upon opening.

Burst disk valve distorted but not penetrated as designed.


Further disassembly showed that the O-ring seal separating the hot and cold gas circuits around the hammer that penetrates the burst disk appeared damaged from heat. That seal damage was allowing the cold gas to escape into the hot gas circuit and then vent. Further, the O-ring prevented hot gas from getting to the subject O-ring around the piston that drives the hammer through the burst disk was in two pieces and showed clear evidence for melting at the edges. Thus, when the dual-redundant initiators fired, the piston O-ring failed (or had previously failed, although it was undamaged when installed) which allowed hot gas to leak past the piston (which nonetheless hit the burst disk hard enough to dent it but not tear it) and to damage the O-ring separating the hot-gas and cold-gas circuits in the valve. These two damaged O-rings then allowed cold gas to vent via the hot gas circuit, resulting in the payload seperating from the rocket.

Naturally, none of these failures ever occured in previous ground testing.

Wind shear was considered as a cause for the sudden change in vehicle direction witnessed during launch right after clearing the rail. Even in calm wind conditions on the ground, there have been past launch events at the MTA which have had sharp unseen discontinuities in the wind profile causing serious perturbation of the flight path in a rocket flight. This potential cause can not be fully excluded, but it is thought to be unlikely..

The venting of the hot and cold gas _may_ have caused the sudden pitch over as seen in video footage. As of now, this is being carried as a working hypothesis.  However, none of this explains why the initiators apparently fired a few fractions of a second after lift-off.

The telemetry data will soon be downloaded from the ground station to see if there was any indication of the beginning of this sequence of events. Because the ground station showed loss of signal (LOS) at 119 feet, and that LOS appears to have been the result of the antenna snapping off in the course of the sudden pitch change. There might not be any recorded data of the relevant accelerations or rates from the ground station.

This report will be updated as new information becomes available.

Examining the launch rail and supporting cables before the planned lowering.
Former RRS member, Kevin Sagis helps in gradually releasing the come-along chain bringing the heavy launch rail back to horizontal as the rest of the team managed the straps.

In conclusion of that day’s launch event, with the recovered parts from the rocket payload examined and packed for shipment back to Bill’s home, the remaining team worked to carefully lower the launch rail back to horizontal using the reversed process used to successfully and safely raise it. The launch rail support legs were left at the MTA as Bill and Mike Pohlmiller were going to consider a new design approach using the same T-slot backbone. Although there was no evidence of the rocket hanging up on any discontinuity, some repairs of the interconnections between the three segments should allow the combined rail path to be more straight.

The RRS is grateful to the many members and participants we had over those several few days. It was a big success despite some significant challenges and disappointment in the results. The project was designed to be a pathfinder to subsequent large solid motor projects and we expect the next motor build and improved payload system design in the new calendar year, 2022.