Failure Knowledge Database / 100 Selected Cases

Failed Launch of H-II Rocket No. 5February 21, 1998 at the Tanegashima Space Center
Masayuki Nakao (Institute of Engineering Innovation, School of Engineering, The University of Tokyo)
National S pace Development A gency of Ja pan (NASDA) l aunched CO METS, the C ommunication an dBroadcasting Engineering Test Satellite by its 5th H-II launch vehicle, from the Tanegashima Space Center onFebruary 21, 1998.
The first firin g of the second-stage en gine followin g the f irst-stage main en gine co mpleted nor mally. The
second burning time period of th e second-stage engine was shorter than schedu led, which was caused by gas
leaking from the combustion chamber of the LE-5A second stage, overheating engine parts and damaging theelectrical system in the engine control box (ECB), resulting in the engine’s premature shutdown.
The H- II launch vehicle r eleased C OMETS at the s cheduled time, but a t a wrong a pogee altitude. W hile t heperigee altitude was close enough to the altitude of the planned release point, the apogee altitude was 1,902 km,a lot lower than the planned altitude of 35,976 km. It failed to place the COMETS satellite into the correct orbit
1. Event
NASDA launched COMETS, the Co mmunication and Broadcasting Engineering Test Satellite with its 5th
H-II launch vehicle, from the Tanegashima Space Center on February 21, 1998.
The first firing of the second-stage engine following the first-stage main en gine completed nor mally, but
the second burning time period of the second-stage engine was shorter than scheduled.
The H-II la unch vehicle released COMETS at the scheduled time, but at a wrong a pogee altitude. While
the perigee altitude was close enough to the altitude of the planned release point, the apogee altitude was1,902 km, a lot lower than the planned altitude of 35,976 km. It failed to place the COMETS sat ellite into
the correct orbit
2. Course
NASDA l aunched COMETS, the Com munication a nd Broadcasting Engi neering Test Satellite by its 5t h
H-II launch vehicle, from the Yoshinobu Launch Site of the Tanegashima S pace Center at 16:55 JST onFebruary 21, 1998. It was launched with the azimuth of 92.5 degrees over the Pacific Ocean (with the true
north at 0 degrees, 92.5 degrees in the clockwise direction pointing almost to the west).
The weather was rainy with a northeast wind of 8.7m/sec. and temperature of 15 degrees C.
The first-stag e main en gine and th e so lid rocket booster fi red n ormally. Th e sol id rocket b ooster, th esatellite f airing a nd the f irst-stage r ocket wer e separated at 96 s econds, 2 43 s econds and 356 seconds
respectively after liftoff.

1Failure Knowledge Database / 100 Selected Cases

Figure 1. Schematic Diagram of NASDA’s H-II Rocket (No. 5) [1]
The fir st firing of the s econd stage engin e began at 362 seconds af ter l iftoff. All in cluding the guidancecontrol went normally as planned until the first firing of the second stage engine stopped 672 seconds afterliftoff.
The second firing of the second-stage engine started 1,410 seconds after liftoff. It was designed to continue
firing until 1,598 seconds after liftoff; however, a series of malfunctions occurred at about 1,450 secondsafter liftoff (40 seconds af ter the start of the second firing). The second firing of the second-stage engine
stopped 1,457 seconds after liftoff (47 seconds after the start of the second firing).
The H- II launc h veh icle s tayed coasting by iner tia for the r emaining firing duration, then released
COMETS near the scheduled time of 1,638 seconds after liftoff.
The C OMETS sate llite w as r eleased a t t he perigee altitude of 246.2km, which w as c lose t o th e pla nned
altitude of 25 0km, however, at th e apogee altitude of 1 ,902 km, which was a lo t lower than the plann ed
altitude of 35 ,976 k m. Despite a su ccessful s eparation fr om th e r ocket, C OMETS fail ed to enter t he2

Failure Knowledge Database / 100 Selected Cases
geostationary transfer orbit.
3. Cause
Figure 2 shows the combustion chamber of the second stage LE-5A engine, and Figure 3 its cross section.
(1) Loss of engine thrust
Hot combustion gas penetrated the weight-saving windows on the LE-5 engine’s nozzle skirt (Figure2). This gas cut the electrical p ower wiring t o the engine c ontrol b ox ( ECB) loc ated near theweight-saving windows on the engine nozzle skirt. The loss of electrical power closed the mail valve
to shutdown the engine.

Figure 2. LE-5A Engine (Exterior) [1](2) Combus tion gas leakage
At about 41 seconds into the second firing of the secon d-stage engine, a crack in the brazin g caused
hot combustion gas to heat up cooling tubes in the lowest part of the combustion chamber. One second
later, the overheated tubes buckled, allowing a large amount of combustion gas to escape through gaps
between th e brazed tub es (0.1 7kg/sec through a 5cm 2 ar ea). As the co mbustion p ressure de creased,liquid hydrogen leak ed fro m the da maged cooling tubes o f th e cooling sy stem (0.04kg/sec). These3

Failure Knowledge Database / 100 Selected Cases
gases heated the ECB electrical power wiring to above 1,500 degrees C. The ele ctrical power supply
was cut off at 46 seconds into the second firing of the second-stage engine.

Figure 3. LE-5A Engine Diagram [1](3) There are two possible explanations to how the brazing failed.
a.

The brazing of the tubes was relatively weak at the center of weight-saving windows where brazed
metal was only present on the inside. In the second captive firing test of the LE-5A engine, cracks
started from the outside and which then developed into openings at the end of the first firing of thesecond-stage engine. Liquid hydrogen started leaking and grew bigger during the second firing of
the second-stage engine. (Probable cause I)

b. Shape irregularity increased at the center of weight-saving windows where cooling tubes were notbrazed to the cas ing. The st ructure could not withstand multiple firings, and the overheated tubes
4Failure Knowledge Database / 100 Selected Cases
buckled dur ing t he second f iring of t he se cond-stage e ngine, a llowing a lar ge a mount ofcombustion gas to leak through gaps between brazed tubes. (Probable cause II)
In either case, variations (voids and fillets) in the brazing material and the br azing metal brittleness at
the temperature of 500 degrees C also contributed to the failure.
4. Immediate Action
NASDA performed telemeter data analysis, fault tr ee analysis and co nfirmation tests to find the cause of
the engine’s premature shutdown. Their findings were as follows:(1) Propellant leak from the combustion chamber caused the engine’s premature shutdown.
(2) Propellant leaked from the damaged cooling system of the combustion chamber.
(3) Propellant penetrated brazing between the nozzle skirts cooling tubes located in the lowest part of thecombustion chamber.
NASDA then performed the following tests to reevaluate brazing in the combustion chamber of the engine
installed on its 7th H-II launch vehicle, which was scheduled to be launched the next.
• X-ra y inspection• Visual inspection: Fillet brazing of cooling tubes in the interior of the combustion chamber casing• Visual inspection: Fillet brazing of cooling tubes around the windows on the nozzle skirt in the exteriorof the combustion chamber casing• X-ray CT inspection of the combustion chamber
Inspections did not find faulty brazing that might have caused propellant leak.
5. Countermeasure
Improvement an d throu gh insp ections w ere made to the brazing and the nozzle structure o f the LE- 7A
engines of the 7th H-II launch vehicle and the LE-5A engines of the H-IIA launch vehicles (the successor
aircraft to the H-II launch vehicle) to ensure the higher reliability of the engines.
A low-pressure firing was reported during the second captive firing test of the LE engine; however, it w as
due to a leak detector jig left in the engine. Turbine propellant was unable to flow into the engine throughthe valve, which was taken up by the jig. It was only the tank pressure that fed propellant to the engine,which resulted in low-pressure combustion. The testing procedures were modified to enforce the placementof a red tag o n an en gine th at has a leak detector installed. A r ed ta g is t o be always stored wit h a l eakdetector.
6. Summary
The H-II launch vehicle is a two-stage rocket powered by liquid-hydrogen/liquid-oxygen engines, capable
5Failure Knowledge Database / 100 Selected Casesof launching a 4-to n class satellite into ge ostationary transfer or bit. It is equipped with two huge solid
rocket boosters (SRB), which along with the vehicle’s main engine provide jump in thrust during the firststage thrust.
The 1st stage LE- 7 eng ine is alarge two-burning cycle engine that bur ns liq uid oxygen and liquidhydrogen. It was developed based on the LE-5 technology.
The 2nd stage LE-5A engine is an improved LE-5 engine whose propellants are liquid hydrogen and liquid
oxygen. It offers higher performance and reliability than the LE-5 en gine. The LE-5A engine can reignite
so that the H-II launch vehicle makes a two-step journey to geosynchronous transfer orbit.
The adv anced S RBs ar e poly butadiene co mposite so lid propellant boosters. Unl ike the solid strap-on
boosters ( SSB) equi pped by the predecessor aircrafts su ch as the H-I launc h veh icle, S RBs h ave
hydraulically steerable nozzles that enable guidance and control of the first stage.
Five earlier H-II launches and missions were all successful. The firings of the 2nd stage LE-5 and LE-5A
engines were completed as s cheduled during the past 14 launches of the H- I and the H-II launch vehicles.
The H-II la unch failure in 19 98 occurred after number of successful launches. Engineers may have been
too confident about their technology. A good example is the careless mistake during the captive firing test.
The second captive firing test of the LE-5A engine was aborted 22.4 seconds after the engine started firing,
because the combustion pressure did not reach the maximum.
Engineers later found that t he low- pressure firing wa s due t o a leak detector still connected to the waist
intake valve of the engine by mistake. Turbine propellant was unable to flow into the engine through thevalve. It was only the tank pressure that fed propellant to the engine, resulting in low-pressure combustion.
It was a s imple mistake th at he ld up t he firing tes t and ultimately caused th e f iring failure of t he engi ne
during the mission. Officials later explained that a leak detector was used to make sure that the combustion
chamber and the nozzle skirt was cooled down after the firing test.
Unevenness (vo ids and f illets) o f the br azing occurre d during th e engine m anufacturing m ay have also
contributed to the brazing failure.
7. Knowledge(1) Compared to machining, brazing process is likely to produce unevenness in fillet size and braze joints.
It is necessary to properly design requirements and parts for brazing to maintain an acceptable quality.
The same applies to welding.
(2) A p erformance test must b e pe rformed w ith an accurate test method and w ithout oversig ht to av oid
producing an inaccurate result. An inadvertent oversight caused the failure of the firing test in this case.
Inaccurate te st methods al so pro duce inaccurate test re sults suc h a s fa ilure a nd pe rformancedegradation.
(3) Continued su ccess in cubates p otential fail ure, be cause it tends t o lead to over confidence and lesssubsequent mindfulness. A development team must tighten its parameters and improve its management,
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Failure Knowledge Database / 100 Selected Cases
especially when it achieved a breakthrough in its project. The same applies corporate management.
8. Background
The H-II l aunch vehicle was manufactured 100% domestically in Japan. Based on th e achievements of its
predecessor aircrafts, it wa s designed to offer low-cost and high reliability for launching of geostationary
satellites and payloads into orbits.
In 1984, NA SDA sta rted the f ull-scale development of the H-II la unch v ehicle and its lau nch site. Aftercompleting conceptual design, system design, basic design and detailed design, NASDA conducted groundsystem tes ts us ing the ground test v ehicle (GTV) for about 6 months fr om Sep tember 19 91 u ntil March
1992.
Before the launch of the 5th vehicle in 1998, the H-II launch vehicle was util ized in a to tal of 5 launches
since the first successful launch in 1994 (No. 1 for experimental missions on February 4, 1994; No. 2 foran experimental mission on August 28, 1994; No. 3 for an experimental and satellite missions on March 18,
1995; No. 4 for a satellite mission on August 17, 1996; and No. 6 for observation and test satellite missions
on November 28, 1997).
While scheduled launch of the No. 5 (COMETS) was postponed due to malfunction of the Advanced EarthObserving Satellite (ADEOS) in 1997, the No. 6 was launched as scheduled. This caused the launch order
to disagree with the vehicle numbering.
References[1] National Space Development Agency of Japan (1998) “Launch Failure of the 5th H-II Launch Vehicle / theCommunications and Broadcasting Engineering Test Satellite (COMETS) “Kakehashi””,
http://www.nasda.go.jp/press/1998/06/comets_980611_01_j.html#1

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