Document
SA396 File 2-0160
AIRCRAFT ACCIDENT REPORT
Adopted: July 17, 1968
u
HONEYWELL, INC.
GRUMMAN G - 159, N861 H NEAR LE CENTER, MINNESOTA JULY 11, 1967 NATIONAL TRANSPORTATION SAFETY BOARD DEPARTMENT OF TRANSPORTATION WASHINGTON D.C. 20591 For sale bg Clearinghouse for Federal SeientiRe and Technical Information, U.S. Department of Commerce, SpringReld. Va. 22151. Annual sub?cription Price $12.00 Domestic; $15.00 Foreign; Single copy $3.00; 311croficlle $O.G5. Order Sunlber PB l i 7 339-9.
NATIONAL TRANSPORTATION SAFETY BOARD D E P A R T M E N T O F TRANSPORTATION WASHINGTON. D . C .
I N C .
HONEYWELL.
G R U " A N G.159. ~ 8 6 1 ~ NERR LE CENTER. MINNESOTA JULY 11. 1967 TABLE O F CONTERPS
. Page
Synopsis . . . . . . . . . . . . . . . . . 1
Probable Cause . . . . . . . . . . . . . . 2
Investigation . . . , . . . . . . . . . . 3
I .
History of the Flight . . . . . . . . . . . 3
1.1
1 . 2 Injuries t o Persons . . . . . . . . . . . 5
Damage t o the Aircraft . . . . . . . . . . 5
1.3 . . . . . . . . . . . . . . .
1.4 Other Damage 5
. . . . . . . . . . . . . 5
Crew Information 1 . 5
1.6 Aircraft Information . . . . . . . . . . . 5
Meteorological Information . . . . . . . 6
1.7
Aids t o Navigation . . . . . . . . . . . . 6
1.8
. . . . . . . . . . . . . . 6
Communications 1.9 1.10 Aerodrome and Ground F a c i l i t i e s 6 . . . . . .
1.11 Flight Recorders 6 . . . . . . . . . . . . .
1.12 . . . . . . . . . . . . . . . . .
Wreckage
1.l2A Pertinent Propeller Information . . . . . . 8
1.12B Examination of Pertinent Engine and
Propeller Controls and Systems . . . . . 10
1.13 F i r e . . . . . . . . , . . . . . . . . . . 11
1.14 Survival Aspects . . . . . . . . . . . . . 11
Tests and Research . . . . . . . . . . . . 11
1.15
1.16 . . . . . . . Other Pertinent Information 13
2 . Analysis and Conclusions 13 . . . . . . . . .
2.1 Analysis . . . . . . . . . . . . . . . . . 13
2.2 Conclusions . . . . . . . . . . . . . . . 19
( a ) Findings . . . . . . . . . . . . . 1 9
. . . . . . . . . .
( b ) Probable Cause 2 1 . . . . . . . . . . . . .
Recommendations 3 .
AppendixA . . . . . . . . . . . . . . .
F i l e No. 2-0160 NATIONAL TRANSPORTATION SAFETY BOARD D E P A R T M X N T OF TRANSPORTATION AIRCRAFT ACCIDENT REFORT Adopted: July 17, 1968 HONEYWELL, INC.
GRUMMAN G-159, ~ 8 6 1 ~ NEAR LE CENTER, M I N N E S O T A m y 11, 1967 SYNOFSIS A t 1034 c.d.t., July 11, 1967, a Grumman G-159, (Gulfstream I) N861H, owned and operated by the Honeywell Corporation, departed from t h e Minneapolis- S t . Paul International Airport, Minneapolis, Minnesota, The purpose of t h e flight was t r a i n i n g on the a i r c r a f t f o r a company p i l o t . The weather was clear.
A t 1100, the f l i g h t while i n communication with the Minneapolis A i r Route Traffic Control Center t o receive VFR radar advisory service, made an emergency request for a radar vector t o the nearest airport. The p i l o t indicated f i r e ~. . . ~ ~ ~ i n one, possibly both engines. J u s t before 1104 he advised, " We're crash r landing. " The radar t a r g e t of N861H was l o s t at 1104.
Between 1100 and 1104, ground witnesses located about 5 miles south- southeast of Le Center, Minnesota, saw the a i r c r a f t approach on e north - easterly heading and begin a wide,right c i r c l e . A s the c i r c l e progressed, white smoke o r vapor was seen coming from the r i g h t engine nacelle Wing area and the l e f t propeller was stopped. Shortly thereafter, the r i g h t propeller stopped, f i r e appeared i n the r i g h t engine nacelle wing area, and about the same time there was an explosion and pieces of wing separated from the a i r c r a f t .
~..
- 2 - The a i r c r a f t immediately went out of control, crashed and burned.
Both p i l o t s , the only occupants of the a i r c r a f t , were f a t a l l y injured.
The a i r c r a f t was destroyed.
The Safety Board determines t h a t the probable cause of t h i s accident was overtemperaturing of both engines, i n f l i g h t f i r e and explosion caused by t h e f a i l u r e of t h e "2" relay i n t h e propeller automatic c r u i s e pitch lock retraction system.
- 3 - 1. INVESTIGATION 1.1 Hisaory of the Flight On the morning of July 11, 1967, a Grumman G-159, (Gulfstream I), ~86111, owned and operated as a corporate a i r c r a f t by Honeywell, Inc., was scheduled f o r an approximate 2 - hour l o c a l V F R , no - flight - plan f l i g h t from the Minneapolis - St. Paul International Airport at Minneapolis, Minnesota.
The purpose of the f l i g h t was t r a i n i n g t o prepare one of the company p i l o t s f o r a G-159 rating. The instructor p i l o t was t h e company's Manager of Flight Operations. The weather was c l e a r throughout the area. Departure was at 1034 w i t h the instructor i n the r i g h t p i l o t s e a t and the t r a i n e e i n the l e f t .
After takeoff, the f l i g h t adv-ised Minneapolis Tower it would be oper - ating 30 t o 40 miles west of the a i r p o r t between 9,000 and 11,000 f e e t , and requested VFR radar advisory service. A t 1058, radar contact was established and the f l i g h t was i n communication with t h e Minneapolis A i r Route Traffic Control Center on frequency 125.9 MHz.
A t 1100, the Center received an emergency c a l l from the f l i g h t requesting a radar vector t o the nearest airport. The a i r c r a f t , observed on radar t o be on a northeast heading, was advised t o t u r n t o 240" f o r t h e Mankato, Minnesota Airport; however, no t u r n was observed. The p i l o t of ~ 8 6 1 ~ then asked
f o r the distance t o Mankato and was advised it was 14 miles and t o
reverse course. A t t h i s time the p i l o t stated, " W e got a dual f i r e -
single f i r e now. " Asked h i s s i t u a t i o n s h o r t l y thereafter, he answered, ". . .
w e had a f i r e i n the l e f t engine - got it out now, we think. " H e again asked
1/ A l l times are c e n t r a l daylight based on the 24 - hour clock.
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for the heading t o Mankato and w a s t o l d 240". Asked again about h i s
s i t u a t i o n the p i l o t answered, " Probably going t o land short - you
b e t t e r get somebody out here. " A t 1103:35, he radioed i n a hurried voice ". . .we're crash landing. " This was t h e last transmission from t h e f l i g h t , and a t 1104 the radar t a r g e t of the a i r c r a f t was l o s t .
During the above sequence of events, ground witnesses near Le Center, Minnesota, which is about 40 miles southwest of t h e Minneapolis - St. Paul International Airport, saw the a i r c r a f t approach on a northeasterly heading and being a wide, right turn. When the a i r c r a f t reached a southerly heading, some witnesses saw a stream of white smoke or vapor t r a i l i n g from t h e r i g h t engine and others noted t h a t the l e f t propeller m s stopped. A s the t u r n progressed t o a westerly heading, the a i r c r a f t descended t o between 300 and 500 f e e t above the surface. The t r a i l i n g smoke o r vapor turned gray or black and t h e right propeller was observed t o slow and stop. While the a i r c r a f t was turning t o t h e northwest, f i r e appeared i n the r i g h t engine nacelle and wing area. Almost simultaneously t h e r e was an explosion i n t h e wing, and pieces separated from the wing area. The a i r c r a f t immediately went out of control and crashed. Ground f i r e consumed major portions of t h e air - c r a f t wreckage.
A computation based on the regular sequence of t r a i n i n g maneuvers for t h e f l i g h t and t h e norma times f o r t h e i r execution indicated that when t h e accident sequence started, t h e p i l o t s were probably engaged i n approaches t o stalls i n one of t h e various a i r c r a f t configurations, or t h e canyon - 5 - approach. A slow airspeed and a rapid application of appreciable power muld be common t o these maneuvers.
1.2 Injuries t o Persons Both p i l o t s , the only occupants of t h e a i r c r a f t , received f a t a l in - juries. There were no i n j u r i e s involving other persons.
Post - mortem pathological and toxicological examinations of the p i l o t s revealed no evidence of a human f a c t o r involvement i n the accident.
1.3 Damage t o the Aircraft Destroyed by impact and f i r e .
1.4 Other W a g e The a i r c r a f t crashed on a cultivated bean f i e l d causing property damage t o the crop.
1.5 Crew Information The instructor and trainee p i l o t s were both qualified and properly certificated f o r the f l i g h t . See Appendix A f o r detailed crew information.
1.6 Aircraft Information N86U wits a G m a n G-159, Gulfstream I. Aircraft records indicated that the a i r c r a f t had been maintained i n accordance with applicable Federal Aviation Administration (FAA) requirements. See Appendix A.
There were no writeups of any r e p e t i t i v e discrepancies i n the a i r c r a f t
f l i g h t logbook. The l a s t writeup A s on July 6, 1967. This indicated that
during takeoff climb a t 175 knots t h e p i l o t s had experienced'a momentary decrease i n r.p.m. and increase i n turbine temperature f o r t h e r i g h t engine.
. - 6 - The writeup indicated t h a t a t the time the propeller cruise pitch lock out l i g h t s were out and the lock was i n place. The writeup indicated there was no apparent reason f o r t h e problem. No corrective action was taken. There had been no other p r i o r or subsequent writeups of the problem.
1.7 Meteorological Information Weather was clear.
1.8 Aids t o Navigation Not involved.
1.9 Communications There were no communications d i f f i c u l t i e s .
1.10 Aerodrome and Ground F a c i l i t i e s Not involved.
1.11 Flight Recorders None was i n s t a l l e d or required on t h e a i r c r a f t .
1.12 Wreckage The a i r c r a f t crashed i n an open cultivated f i e l d approximately 5 miles
southeast of Le Center, Minnesota. - It struck t h e ground r i g h t wing down
and disintegrated along a northwest ground path, 1,220 f e e t long and 264 f e e t wide. A t impact,the landing gear and f l a p s were up and t h e propellers of both engines were feathered.
A section of the r i g h t wing lower skin from between wing s t a t i o n s 164 and 293, which includes portions of the right wing f u e l tank, wheelwel1,and engine nacelle areas, was found 930 f e e t from the i n i t i a l impact point back
- 2 / Geographical location: Latitude N40°-20' Longitude W93"-44'
- 7 - along the f i n a l flightpath. This structure and other pieces of the wing and pieces of engine nacelle found a t o r near t h i s location showed evidence of intense i n f l i g h t f i r e and t h e force of an explosion.
Along the ground from the major wing pieces t o the i n i t i a l impact point there were numerous burned and molten pieces from the r i g h t wing engine nacelle area. Smaller b i t s and pieces of burned and molten metal were a l s o located a s far back a s 3,200 f e e t along the f i n a l flightpath. The r i g h t main landing gear Was recovered outside the ground f i r e zone. The t i r e s and other components showed evidence of exposure t o intense inflight f i r e .
i n the main wreckage area, structure from locations adjacent t o t h e separated wing and nacelle pieces showed evidence of i n f l i g h t f i r e . The f i r e patterns indicated t h a t the f i r e was concentrated i n t h e r i g h t engine nacelle area. The structure also showed evidence indicating t h a t explosion forces were generated i n the wing tank.area. The balance of the airframe wreckage located i n the main wreckage area revealed no additional evidence significant t o the accident.
Examination of t h e powerplants of the a i r c r a f t disclosed t h a t during the ground impact sequence, the l e f t propeller remained with i t s engine while the right one was broken away. Both propellers were feathered when i n i t i a l impact occurred.
Both engines had been subjected t o extreme and destructive pre - impact operating overtemperature which was concentrated i n t h e turbine sections and near the tops of the engines i n the combustion chamber areas. In t h e
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r i g h t engine, t h e r e were indications of i n f l i g h t fire i n t h e t o p section of t h e a i r c r a f t f i r e w a l l i n t h e engine mount area. Among numerous other indications of engine overtemperature, t h e major portions of most i n t e r - mediate and high - pressure turbine blades were melted away. These blades a r e made of Nimonic ( n i c k e l a l l o y ) material, capable of withstanding approximately 2,400°F., compared t o t h e normal maximum operating engine turbine temperature of about l,500°F.
1.12A Pertinent Propeller Information The Dowty-Roto1 propeller system as i n s t a l l e d on t h e a i r c r a f t includes a s e t of propeller c r u i s e p i t c h locks on each propeller. These are designed t o prevent t h e propeller blades from f i n i n g off (decreasing) below 34.5 " i n t h e event of a malfunction or f a i l u r e at a high airspeed which could otherwise r e s u l t i n a dangerous propeller overspeed.
I n normal operation of t h e system there i s an automatic e l e c t r i c a l cruise pitch lock withdrawal feature t o r e t r a c t t h e locks when t h e propeller blades f i n e off aerodynamically during a decrease i n airspeed. I n an airspeed range of 160 t o 175 knots, propeller blade angles reach 36.5", at which point a propeller hub switch i n each propeller closes, permitting e l e c t r i c a l current t o flow t o t h e " X " r e l a y ( l e f t engine) and t o t h e "Z" r e l a y ( r i g h t engine).
When both relays receive e l e c t r i c a l power and t h e i r e l e c t r i c a l contacts close, current flows t o t h e f l i g h t s a f e t y lock switch and then t o t h e l e f t and r i g h t propeller c r u i s e p i t c h lock removal solenoids. When t h e solenoids are ener - gized, they operate t o cause engine o i l pressure t o extract the c r u i s e locks, allowing t h e propeller blades t o f i n e off so t h a t appropriately high r.p.m.
may be a t t a i n a b l e at t h e lower airspeeds.
- 9 - A pair of lights, the cruise pitch l i g h t s on the p i l o t ' s instrument pne1,are also operated by the propeller hub switches and a r e illuminated i n unison when e i t h e r propeller i s a t 36.5 " or below. A second p a i r of lights, the cruise lock out lights,located below the cruise p i t c h l i g h t s , are illuminated separately by the application of engine o i l pressure t o remove the cruise pitch locks. Both the " X " r e l a y and the "2" r e l a y must ,function f o r these l i g h t s t o come on and f o r the c i r c u i t t o be completed through the f l i g h t safety lock switch.
The propeller system design incorporates a provision f o r circumventing the automatic cruise pitch lock withdrawal system i n case of a malfunction or failure which might remove the locks prematurely or remove them when they need t o remain i n place, such as during a propeller overspeed. Placing t h e flight safety lock switch t o emergency position breaks the e l e c t r i c a l c i r c u i t at the switch and prevents automatic cruise pitch lock extraction and the propeller blade angles cannot decrease below 34.5 O , The system also incorporates a provision t o extract the locks manually. Positioning of t h e high pressure f u e l cocks, f u e l and propeller control levers on the power quadrant, t o the cruise lockout position w i l l , regardless of t h e condition of the automatic removal system and position of i t s switches o r the availa - b i l i t y of e l e c t r i c a l power, open separate valves mechanically and port engine o i l pressure t o the extraction side of the cruise p i t c h locks t o extract them.
In the event the cruise pitch locks do not extract properly, t h e warning t o the p i l o t would be f a i l u r e of the cruise p i t c h lockout l i g h t s t o come on.
If the cruise pitch locks were not extracted at low airspeed, and an attempt were made t o apply power, the turbine sections of the engines would be subjected t o overtemperature. T h i s i s because the Rolls Royce Dart engine control system i s correlated with the propeller control system i n such a way t h a t the engine control system depends on the propeller control system for governing airflow a t given f l i g h t and power lever conditions. Thus, under normal conditions,acceleration f u e l under a power application would be matched with a large increase i n airflow through rapid engine spool - up. I f , however, the engine were i n an overloaded condition because of propeller blade hang - up a t the cruise pitch lock angle and could not accelerate fast enough t o alleviate very high f u e l - a i r ratios, it would be subjected t o a rapid and extreme over- temperature.
1.12B Examination of Pertinent Engine and Propeller Controls and Systems Examination of cockpit engine and propeller controls revealed t h a t both power levers were at or near idle. The high - pressure cocks were f u l l rearward, the propeller feather position. The ground f i n e lever was i n the f l i g h t position.
The l e f t engine f i r e extinguisher T - handle was pulled out and the f i r e extinguisher b o t t l e switch was positioned t o the No. 2 b o t t l e . The r i g h t engine f i r e extinguisher T - handle was broken off but the mechanism showed it had been i n the stowed position when it was broken off. Both engine f i r e extinguisher b o t t l e s were empty. The l e f t engine f u e l shutoff valve was closed; the r i g h t was open.
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The f l i g h t safety lock switch, a rocker type, was found. i n the emer - gency position. The u n i t and surrounding structure, however, had received considerable impact damage. The top mounting screw was i n t a c t , but the lower mounting screw was t o r n out and the lower end of the switch was pulled out of the panel about one - quarter inch. The p l a s t i c cover of the rocker element of the switch was knocked o f f and the panel molding around the switch was twisted and broken.
The " X " and "Z" relays and other pertinent components of the propeller operating system hereinbefore described were recovered. These, a s well as fuel system components and key accessories, were checked functionally and/or internally examined.
Fuel, o i l and water - methanol samples were tested, and an indentification analysis was made of samples of metal s p a t t e r found i n the l e f t engine
exhaust unit. -
1.13 Fire
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T h i s accident involved a f i r e and explosion i n the right wing engine nacelle areas p r i o r t o impact.
1.14 Survival Aspects The accident was nonsurvivable.
1.15 Tests and Research Fuel, o i l and water - methanol samples given laboratory examination were found t o be within specifications and without contamination.
A sample of metal s p a t t e r found i n the l e f t engine exhaust u n i t was analyzed and found t o be composed of the same material a s the turbine blades.
3/ See Section 1.15, Tests and Research.
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The f u e l pumps, f u e l control units, propeller control u n i t s and low torque switches from both engines were checked. A l l were capable of normal operation except t h e r i g h t engine f u e l pump and f u e l control u n i t which were too badly damaged f o r functional checking. Teardown inspection of these components, however, showed no evidence of pre - impact malfunction o r f a i l u r e .
The f l i g h t s a f e t y lock switch, the propeller pitch lock u n i t s and hub contact switches were checked and found t o operate normally.
The propeller junction box containing the " X " and "2" relays was re - covered, and the relay u n i t s were given laboratory examination and functional checks. The " X " relay showed evidence of considerable arcing and material t r a n s f e r on its contact points, but it operated normally. When the " Z " relay was tested,the normally closed contacts remained open i n t h e power - off con - dition. When the u n i t covers were then sawed o f f , i t was found t h a t three of the four contact leaves were broken and t h e contacts showed minimal arcing and material transfer. The f a i l u r e s of the leaves were i n fatigue. A voltage check revealed t h a t t h e voltage necessary t o make the r e l a y operate varied from a minimum of 24 t o 40 volts, with most checks requiring well above 24 volts. The a i r c r a f t e l e c t r i c a l system i s a 24 - volt system.
From the r e s u l t s of the examination and t e s t i n g , it was considered t h a t the " X " relay possibly would have been intermittent i n operation, and, at best, the operation of t h e "2" relay would have been unreliable.
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1.16 Other Pertinent Information A review of the Grt~mman approved training manual and the Honeywell training data revealed t h a t they explained the operation of the automatic cruise pitch lock system and related panel indicator l i g h t s . They f u r t h e r explained the use of the f l i g h t safety lock switch, i t s effect on the auto- m t i c cruise pitch lock withdrawal system, and the manual cruise pitch lock extraction procedure using the high - pressure f u e l cocks.
The Gnunman and Honeywell f l i g h t manual procedures require positioning the high pressure f u e l cocks i n the cruise pitch lockout position before and during takeoff and before, during and a f t e r landing. They do not, however, c a l l f o r such positioning f o r low airspeed f l i g h t o r f o r maneuvers involving a combination of low airspeed and high power recovery.
2. ANALYSIS AND CONCLUSIONS 2.1 Analysis The physical condition of the engines of N86lH showed that both had been subjected t o severe overtemperature during operation. While the damage t o the l e f t was more severe than t o the right, it was the same kind, and i n both instances required i n f l i g h t engine shut down.
The severe overtemperature damage t o the turbine sections of the engines could only have resulted from an excessive amount of f u e l being introduced f o r the engine operating conditions. Such an over - rich mixture could have resulted from one of two causes: e i t h e r from a grossly malfunctioning f u e l c o n t r o l unit of each engine occurring a t the same time, o r from an engine r.p.m. and resultant mass airflow too low f o r normally scheduled fuel.
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Examination and t e s t s of the f u e l control u n i t of each engine eliminated t h e first of these possible causes for over - richness. This leaves only the inordinately low r.p.m. which would have been the r e s u l t of propeller blade hang - up of both propellers on the cruise pitch locks upon power lever advance - ment at low airspeed. A s previously explained, an attempt t o increase power under these condtions would r e s u l t i n an especially rapid and high rise i n engine gas - turbine temperature. Normally, acceleration f u e l would be matched by a large increase i n airflow through fast engine spool - up. However, with the engines overloaded due t o the combination of high propeller blade angle and low airspeed, acceleration could not occw f a s t enough t o alleviate t h e extremely high fuel - air r a t i o s and consequent overtemperaturing of the engines.
A rapid high power application which would be used i n recovering from a practice approach t o a stall, o r during the l a t t e r phase of a canyon'approach, would complete the conditions f o r very high engine temperatures. In t h i s instance, as evidenced by the melted turbine blades, the temperature i n the engine turbine areas m s a t l e a s t 2,400"F.
I n t h i s accident,the propeller blades could have hung up on the cruise pitch locks f o r one of t w o basic reasons. The first i s t h a t the f l i g h t safety lock switch was positioned t o emergency. A s previously described, with the switch i n t h i s position the automatic cruise pitch lock removal system would be inoperative. I n such an event, i f airspeed were reduced and t h e cruise pitch locks were not withdrawn manually by positioning the f u e l cocks t o cruise lockout, t h e propeller blades would hang up on t h e locks creating the
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overloaded engine condition. The second reason i s t h a t a malfunction o r failure of e i t h e r o r both of the “X” and “2” propeller e l e c t r i c a l system relays occurred, disabling the automatic cruise p i t c h lock removal system and thereby creating the same overloaded engine condition.
Under the circumstances of e i t h e r of these reasons,the cruise lockout l i g h t s would not have come on indicating withdrawal of the locks. Thus, under e i t h e r situation it must be presumed t h a t the p i l o t s f a i l e d t o note that the l i g h t s did not come on and did not take the necessary remedial action t o remove the locks manually with the high - pressure f u e l cocks. Had t h i s action been a matter of f l i g h t procedure f o r reduced - speed maneuvers, as called f o r during landing and takeoff, it presumably would have been done, and the consequences of the overloaded engine s i t u a t i o n r e s u l t i n g from e i t h e r cause would have been averted. The Safety Board concludes t h a t t h e absence of t h i s f l i g h t procedure was a deficiency i n t h e FAA - Approved Flight Manual
4/
f o r the a i r c r a f t .
Several explanations f o r the f l i g h t s a f e t y switch t o have been i n the emergency position were considered but,in doing s o , l i t t l e weight was attached t o the f a c t t h a t t h e switch was found positioned t o emergency.
This i s because damage t o the switch and surrounding panel structure indi - cated that impact could have accounted f o r t h e position.
It was considered t h a t , since t h e f l i g h t was f o r training, the in - structor p i l o t may have moved the f l i g h t safety lock switch t o its emergency position t o t e s t the trainee’s reaction. This p o s s i b i l i t y i s considered See Recommendations, Section 3.
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remote because the t r a i n i n g syllabus did not c a l l f o r such check,and t h e f l i g h t was a t r a i n i n g rather than a check f l i g h t . Moreover, i n view of the consequence of leaving the switch i n t h i s position, it i s most d i f f i c u l t t o imagine the i n s t r u c t o r having positioned it t o emergency and not taking t h e necessary corrective action if t h e trainee f a i l e d t o do so.
Inadvertent actuation of the f l i g h t s a f e t y lock switch t o emergency was considered a s another p o s s i b i l i t y . The switch on N861H was not guarded and,being a rocker type, only a push on the top portion would be necessary t o actuate it. Despite these factors, t h i s p o s s i b i l i t y is unlikely because the switch was located on the eyebrow panel and it is not next t o other switches o r controls used i n normal f l i g h t .
Another p o s s i b i l i t y was t h a t the switch was positioned t o emergency i n response t o a propeller overspeed condition. This p o s s i b i l i t y was d i s - counted because there i s no evidence of propeller overspeed o r of a f l i g h t s i t u a t i o n conducive t o a propeller overspeed.
Probably the most substantive reason f o r the f l i g h t s a f e t y lock switch t o have been positioned t o emergency would be an abnormal operation of t h e cruise pitch l i g h t s indicating a malfunction of the cruise pitch lock system.
For t h i s t o occur, however, it would require a f a i l u r e of e i t h e r t h e " X " o r the "2" relay i n the system and,simultaneously, an intermittent operation Of probable because of the apparent operabie condition of the " X " relay, and since only one relay i s necessary t o the function of the cruise p i t c h l i g h t s , the p o s s i b i l i t y of e r r a t i c operation of the cruise pitch l i g h t s is discounted.
There are two additional factors which reduce substantially the l i k e l i - hood of any of the p o s s i b i l i t i e s which presume an intentional p i l o t actuation
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of the f l i g h t safety switch t o emergency. F i r s t , it is most improbable that the p i l o t s would continue a t r a i n i n g exercise a f t e r experiencing a situation which prompted actuation of the f l i g h t safety lock switch t o emergency. Secondly, it is equally improbable t h a t they would knowingly move the switch t o emergency and not take the companion followup action of positioning the high - pressure f u e l cocks t o cruise lockout.
I n view of the foregoing, the Safety Board concludes that the f l i g h t safety lock switch was positioned t o emergency by impact and not by i n - tentional o r inadvertent crew actuation. Accordingly, it f u r t h e r concludes that the cruise pitch locks remained i n place because of a f a i l u r e within the automatic withdrawal system. I n t h e i n s t a n t case, t h e only discrepancy found which could have disabled the system was the deteriorated condition of the " X " and "Z" e l e c t r i c a l relays. The physical condition of the relays indicated t h a t e i t h e r o r both could have f a i l e d t o function; however, t e s t i n g of the units indicated that it was most probably only the "Z" r e l a y that failed. The f a i l u r e of one is s u f f i c i e n t t o disable t h e system.
Brplanations of why t h e p i l o t s f a i l e d t o notice t h a t the cruise pitch lockout l i g h t s were not on and remove the cruise pitch locks with the high - pressure f u e l cock, o r t o notice t h e extreme turbine gas temperatures reflected by the turbine temperature instruments when power wasapplied, remain matters of conjecture. It is evident, however, that while engaged i n the power application phase of e i t h e r a stall or a canyon approach, t h e attention of the p i l o t s would have been divided and not directed t o the - 18 - l i g h t s o r power indicators any more o r l e s s than t o f l i g h t instruments.
The negative aspect of the l i g h t indication would a l s o be easy t o overlook.
It i s also noted t h a t under a rapid and substantial power application t h a t would characterize the use of power i n e i t h e r maneuver, the extreme engine temperature and resultant damage would occur very rapidly.
The sequence i n which the engines were overtemperatured i s not clear.
The use of both f i r e b o t t l e s i n the l e f t nacelle and the greater turbine overtemperature damage i n the l e f t engine, although not operated a s long, suggest t h a t it was the first t o overheat. A t the same time there was evidence of greater overheat damage external t o the basic engine i n the i n - stance of the r i g h t engine. The f a c t t h a t the l a t t e r was last t o be shut down would not seem t o have p a r t i c u l a r significance t o t h e sequence question, because a f t e r the l e f t engine was stopped there would be a natural reluctance t o shut down the l a s t available power source.
Physical evidence showed, w i t h a high degree of certainty, t h a t an explosion occurred i n t h e r i g h t wing f u e l tank. It is also evident from the burned right main gear t i r e s there was an i n f l i g h t f i r e i n the r i g h t wheelwell and the' f i r e was caused by overheat of the right engine. It was n o t possible, however, t o determine the a c t u a l f i r e propagation from the engine t o the wheelwell o r whether there was a f u e l leak i n the wheelwell before the explosion. The best explanation seems t o be t h a t the jet pipe was ruptured by the " wash " of f a i l e d turbine blades permitting extreme heat t o enter the wheelwell. T h i s could have been s u f f i c i e n t t o have induced the
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fuel tank explosion. However, i n view of the magnitude of the wheelwell f i r e damage,it i s believed there was an a c t u a l release of f u e l within the -wheelwell i t s e l f which could have been caused by turbine blade " shrapnel " - damage. Then, the combination of superheated j e t flow and actual fueled f i r e would well explain the culminating f u e l tank explosion.
2.2 Conclusions ( a ) Findings 1. The p i l o t s were properly c e r t i f i c a t e d and qualified f o r the f l i g h t .
The f l i g h t was i n i t i a t e d as a t r a i n i n g f l i g h t t o 2 .
prepare the trainee - pilot f o r a type r a t i n g i n the a i r c r a f t .
The f l i g h t progressed without incident u n t i l the time 3.
the performance of stall or canyon approach maneuvers was called f o r i n the t r a i n i n g sequence.
4 . During the power application phase of a s t a l l o r a canyon approach, both engines received destructive overtemperature damage.
The l e f t engine was shut down,followed by shutdown of 5.
the r i g h t engine a f t e r emergency f i r e procedures were executed with respect t o the l e f t engine.
6. Inflight f i r e associated with overtemperature of the r i g h t engine caused an explosion of the r i g h t wing f u e l tank. The explosion damage made the a i r c r a f t uncontrollable.
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7. Overtemperature of the engines during t h e power application resulted from an overloaded condition of both engines due t o propeller hang - up on the cruise pitch locks.
a. The automatic cruise p i t c h lock withdrawal system was disabled
due t o the f a i l u r e of t h e "2" relay and possibly t h e " X " relay i n the e l e c t r i c a l c i r c u i t of the system.
The f l i g h t safety switch was i n the normal position when t h e 9.
engine overtemperaturing occurred and was moved t o emergency position by impact.
10. Before the power application, the p i l o t s did not detect t h a t the cruise pitch lockout l i g h t s were not on, indicating t h e cruise p i t c h locks were not withdrawn. During t h e power appli - cation they did not note the engine temperature indications i n time t o prevent the overtemperature damage.
Engine overtemperature damage occurred very rapidly at low 11.
airspeed under the conditions of high - power f u e l lever s e t t i n g and propeller blade angle too high f o r low airspeed.
The cruise p i t c h locks a r e removed manually by positioning the 12.
high pressure f u e l cocks t o cruise p i t c h lockout. This po - sitioning is a procedure called f o r i n the FAA - Approved Flight Manual f o r t h e a i r c r a f t f o r landing and takeoffs but not f o r slow f l i g h t maneuvers.
The FAA - Approved Aircraft Flight Manual was deficient i n not requiring positioning of' t h e high - pressure fuel cocks t o c r u i s e pitch lockout f o r low airspeed f l i g h t maneuvers.
4 (b) Probable Cause The Safety Board determines t h a t t h e probable cause of t h i s accident was overtemperaturing of both engines, i n f l i g h t f i r e and explosion pitch lock retraction system.
3. R E O M M E N D A T I O N S As a r e s u l t of t h i s accident the National Transportation Safety Board made twu basic recommendations t o the Federal Aviation Administration t o prevent the occurrence of another accident f o r the same o r similar reasons.
The first recommended t h a t consideration be given t o requiring t h e i n s t a l - lation of a flashing red l i g h t on t h e eyebrow panel of G-159 a i r c r a f t which would be activated, if f o r any reason the f l i g h t s a f e t y switch were i n t h e emergency position. It was recommended that a placard a l s o be i n s t a l l e d warning t h a t with the f l i g h t s a f e t y switch positioned i n emergency, the cruise pitch locks must be removed manually when airspeed is reduced below cruise. The red l i g h t would deactivate when the locks were removed manually.
The second recommendation was that a new instruction be incorporated i n t h e G-159 Airplane Flight Manual which would prescribe t h a t the high - pressure f u e l cocks be moved t o t h e cruise lockout position during low airspeed maneuvers, the same a s specified f o r landings and takeoffs.
I n response t o t h e first recommendation, t h e F A A indicated that the installation of the flashing red l i g h t would involve a modification of an extensive and complex nature, and therefore it would be necessary t o develop an a l t e r n a t e course of action t o improve t h e r e l i a b i l i t y and operational s a f e t y of t h e propeller system i n t h e problem area. Accordingly, the Administration took action t o reduce the replacement time of the " X " and "Z" relays from 2,500 senrice hours o r 5 years t o 1,000 service hours o r 12 months. It also required and approved a change t o t h e emergency pro - cedures section of the G-159 r e l a t i n g t o t h e f l i g h t s a f e t y switch. To emphasize t h e importance of following approved operational procedure, a warning note i s incorporated i n the revision t o indicate t h a t engine turbine overtemperature can occur if t h e procedures a r e not followed.
The F A A agreed w i t h the second recommendation, and the G-159 Airplane Flight Manual was revised t o incorporate the recommended f l i g h t procedure.
BY THE NATIONAL TRANSPORTATION SAFETY BOARD: JOSEPH J. O'CONNELL, Jr.
I S I Chairman O S C A R M. LAUREL /SI Member JOHN H. REED I s 1 Member LOUIS M. THAYER / S I Member FRANCIS H. M c A D A M S / S I Member f b m t ion ,atmetor p i l o t , Captain Thomas U. Grove, age 46, held a i r l i n e rt p i l o t c e r t i f i c a t e No. 1469 - 40 with E - 3, CV-340/440 and Grwnman .aircraft ratings and commercial privileges SES, SEL, AMEL and AMES.
-159 rating was acquired March 16, 1965. Captain Grove held a first - @ medical c e r t i f i c a t e issued with no limitations on December 12, 1966.
aptain Grove had accumulated 9,388 t o t a l p i l o t hours, of which 639 the G-159. H e had flown about 3 hours i n the 24 - hour period before ;"&ccident and h i s r e s t time had been 15 hours and 30 minutes.
. .....
.
aptain Grove was employed by Honeywell, Inc., i n 1946. During h i s n p gkbyment he held positions of Flight Test Engineer, Engineering P i l o t . .
I M a g e r of Flight Operations. The l a t t e r position he held f o r 14 years.
j .
1 .
- w s a l s o a registered Professional Engineer i n Minnesota.
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>,,._ . ., $ 3 ' ' I ' Trainee - pilot, Copilot James R. Bradford, age 45, held a i r l i n e trans - . .
P . , .
@@rt p i l o t c e r t i f i c a t e No. 478587 with E - 3 and CV-340/440 a i r c r a f t ratings ;. $>, ; I
$md commercial privileges A I @ & , SEL, and rotorcraft helicopters. He was
signated by the F A A as an Engineering Representative (Flight Test) f o r . .
. 1 :.:@Lectronic f l i g h t control systems. H e held a f i r s t - c l a s s medical c e r t i f i - 2 : .,pate issued with no limitations on March 16, 1967.
P i l o t Bradford had accumulated approximately 9,125 p i l o t hours, of 2.
"$hich 67 were a s copilot i n t h e G-159. H e had flown about 3 hours i n the ..
:. &-hour period preceding the accident and h i s r e s t period had been 15 hours > ' : and 30 minutes.
F'ilot Bradford was employed by Honeywell, Inc., i n 1953. During h i s employment he had held positions of Aircraft Mgineer, Flight Test Engineer, Engineering P i l o t and Project F'ilot/Engineer. For the most recent 9 years he was Project P i l o t and Engineer on I n e r t i a l Guidance, auto - landing, radar altimetry and other company projects. H e ms a graduate E l e c t r i c a l Engineer.
Aircraft Information The a i r c r a f t was a Grumman model G-159 (Gulfstream I), ~ 8 6 1 ~ , inanu- factured with an Airworthiness. C e r t i f i c a t e dated January 5, 1965. Since new t h e a i r c r a f t had accumulated 1937 t o t a l hours including 92 since t h e last major inspection. . ..
> 1 ... : .~ > x '.
*.., The a i r c r a f t was equipped w i & 2 Rolls Royce R - W - 7, model 5 2 9 - a 5- . .
engines,each of which had accumulated 1,937 hours since new and had not .. .'
been overhauled. The engines were equipped with Doxy-Rotol R-184/4-30-4/50 propellers,each of which had accumulated 1 9 3 7 hours since new and had not hih .. : ; +..* i been overhauled.