Document
NASA Tech ni c a l M emor a ndum 80235
NASA-TM-80235 19800012808
Emergency In- F light Egress Opening
f o r Gener a l Avi a tion Aircr a ft
APRIL 19 8 0 r , - _4>- ' J I
APR 1 . _ I_._ . .,_
LANGLEY RESEA R CHCENTE R LIBRARY , NASA J'IAM P -TO_NL V I RGINIA
NASA Technical M emorandum 8023 5
Emergency In- F light Egress Opening
for Gener a l Avi a tion Aircr a ft
La urence J. Bement Langley Re search Center Hampto n , Virgi n ia N ational Aer o nautics and Space Administration Scienti f ic and Techn i cal I nfo r ma t ion O ff ice 19 8 0 SUMMARY I n s u pp o r t o f a s t all / sp i n research pr o gram a t t he Nat io nal Aeronautics and Space Administration (NASA), Langley Research Center (LaRC), an emergency in-flight egress sys t em is being installed in a light general aviation air- plane. T he airplane has no provision for egress on the lef t side. A lef t -side egress opening would greatly enhance the pilo t 's ability for bailout, particu- larly in a right spin. To avoid a major structural redesign for a me c hanical door, an add-on ] ].2-kg (24.6-ib) pyrotechnic-a c tuated sys t em was developed to create an opening in the existing structure. The skin of the airplane will be explosively severed around the side window, across a central stringer, and down to the floor, creating an opening of approximately 76 by 76 cm (30 by 30 in.).
The severed panel will be jettisoned at an initial veloci t y of approximately 13.7 m / sec (45 ft / sec). Sys t em development included a total of 68 explosive severance tests on aluminum material using small samples, small and full-scale flat-panel aircraft structural mock-ups, and an actual aircraft fuselage.
T hese tests proved explosive sizing / severance margins, explosive initiation, explosive product containment, and system dynamics. T his technology is applicable to any aircraft of similar construction.
INTRODUCTION Airplanes, upon stalling, may begin a rotating, sinking motion called a spin. Stall / spin is a prime causal factor in fatal general aviation accidents.
Several light airplan e s are being spin tested at NASA-LaRC in an effor t to improve the stall / spin characteristics of this class of airplanes. T hese air- planes are equipped with tail-mounted spin recovery parachute systems in the event that the spinning cannot be stopped by the normal airplane controls. If both the airplane controls and the recovery parachute fail to stop the spin, the pilot would have to abandon the airplane. One airplane currently being readied for spin testing has a single door on the right side with no option for egress on the lef t side. Bailout would require the pilot to move across the aircraft to open the existing door, possibly against centrifugal loads; this is a difficult task at best. A left-side egress opening would minimize the pilot's bailout effort and ti m e.
A pyrotechnic-actuated egress opening was developed because it proved to be more advantageous than a mechanical system on the basis of structural mod- ification, performance, and the po t ential for success. A mechanical system would require considerable structural modification and reanalysis to incorpo- ra t e a door and release mechanisms. A pyrotechnic approach would be an add- on system, based on previous experience gained in the F- ] ] ] and B-I escape modules.
The pyrotechnic sys t em developed in this effort uses a small-quantity, fully contained, explosive-shaped charge to sever and jettison a left-side panel from the airplane. T he system is ini t iated mechanically by a bell c rank pulled by the pilot. From that point, the system functions automatically.
The design and development capitalized on existing pyrotechnic technology, materials, and components, and emphasized proving all aspects of functional performance. This proof of performance was accomplished analytically and functionally to show m argins of capability greater than the force, strength, or energy required.
This paper describes the design, development, and functional testing of the pyrotechnic-actuated emergency in-flight egress opening for a NASA-LaRC general aviation research airplane.
Identification of commercial products in this report is used to adequately describe the model. The identification of these commercial products does not constitute official endorsement, expressed or implied, of such products or man- ufacturers by the National Aeronautics and Space Administration.
APPARATUS This section describes the off-the-shelf components that were qualified under previous aerospace programs and the test fixtures used in this program.
Flexible Linear-Shaped Charge (FLSC) FLSC has been widely applied by the aerospace community in such applica- tions as rocket vehicle staging and aircraft escape systems. Figure I shows a transverse cross section of FLSC. The materials used in this application are hexanitrostilbene II (organic-precipitated H NS If) explosive (3.]9 g / m (15 grains / ft)) in a silver sheath. This explosive can only be initiated by a high explosive input (greater than 5.5 × 109 N / m 2 (800 000 psi)); it cannot be initiated by gunfire, lightning, electromagnetic-induced radiation, or phys- ical handling. It will burn in a fire, but cannot achieve its cutting func- tion. On initiation, the material detonates at a linear velocity of 7250 m / sec (23 800 ft / sec), generating 2.0 x ]0 ]0 N / m 2 (3 × ]06 psi). The expanding gases and sheath materials are focused in the chevron to effect a metal penetration and deformation / breaking action.
F LS C B oo s t er T i p s To assure reliable initia t ion of t he FLSC and t o seal the exposed explo- sive a t the ends of the six lengths of FLSC, boos t er tips were installed. These tips are cups ( 4.83 mm (0.]90 in. ) i.d. , 8.89 mm ( 0.35 in.) height) stamped from 0.]5-mm (0.006-in.) 302 stainless s t eel (condition A) and loaded wi t h hexani t rostilbene I ( H NS I) at 2. 2 0 × 108 N / m 2 (32 000 psi) t o a height of 3.81 mm (0.]50 in.). T he cups were potted on the ends of the FLSC with a non- solvent structural adhesive (Scotch-Weld Brand Structural Adhesive 22 ] 6 B / A]).
]Produc t of 3M Company.
L an y ard - A c tuated D et o na to r The de t onator used to init i a t e the FLSC is shown in figure 2. A ].27-cm (0.5-in.) stroke compresses the linear spring to 89 N (20 ibf) resistan c e a t release of the sear. T he firing pin asse m bly is driven into t he percussion primer t o initiate the lead azide / HNS I explosive materials in the ou t put cup.
T he lead azide provides an in t erface to develop t he initiation flame to a detona t ion within a 2 .54- m m (0. I -in.) column to properly initiate the FINS I, whi c h in turn ini t iates an FLSC booster tip.
Manifolds To properly locate, secure, and protec t the FLS C boos t er tips, four 6061- T 6 aluminum manifolds were attached to the skin and aircraft s t ructures, one above and below the cen t ral stringer a t the forward and af t ex t re m i ti es of t he egress area. The af t manifolds con t ained a close-tolerance groove t o sec u re the t ips. H owever, the forward manifolds contained no t only a gr o ove but also a threaded por t to re c eive the lanyard-actua t ed de t ona t ors a t the correc t rela t ive position to the tlps (minimum gaps of 0.50 mm (0.020 in.)) to assure reliable explosive ini t ia t ion.
I nternal Containment Development Fixture T o develop the internal structure required to contain the explosive blast of t he FLSC, a t es t fixture was develo p ed that would demons t ra t e performance margins. The fixture was an exa c t mo c k-up of a t ypical a i r c raf t s t ru c ture, bu t t o demonstra t e a containment margin, t he explosive load was increased to 150 percen t of the required a mount and the mock-up of the air c raf t skin thick- ness was in c reased to preven t any severance and venting of the explosive pres- sure. Fur t her, the in t ernal free volume within the containment structure, as well as t he c learances of the FLSC t o the struc t ure, were red u ced to the mini- mums expec t ed in the air c raf t .
Small-Panel Test Fixture Wood-framed pane l s, measuring 45.7 by 45.7 cm (18 by 18 in.) were used as mock-ups of aircraft-representative structures for explosive severance tests.
The mock-up skin was a t tached to the frame, and the representa t ive full-scale structural elements, with explosive components, were mounted to the skin.
Full-scale r flat- panel test fixtures .- T wo full-scale, light-airplane structures were mocked-up in wood-frame d , flat-panels to evaluate the egress system performance. All materials (Alclad 20 2 4-T4), material th i cknesses (1.0 2 mm (0.04 in.)), an d structural layou t s, inclu d ing a 3.175-mm (0.125-in.)
t hick plexiglass window, were mocked-up. A 3.8-cm (l.5-in.) square-mesh s t ain- less wi r e (0.5 mm (0.020 in.) t hick) was used on t he second t es t to prevent the p lexiglass from moving in t ernally. A complete assembly, ex c ept for the initia- tion system , was tes t ed.
Aircraft fuselage test fixture.- A center section of a typical airplane fuselage was modified to si m ulate the end application research-airplane struc- ture as closely as possible, as shown in figure 3.
The research-airplane structure is fabricated from ].02- m m (0.04-in.)
Alclad 2024-T4 aluminum. The fuselage skin panels are made up of flat stock, overlapping above and below the window and just above the floor at the stringers. The depth of the formed channel frames, stringers, and ribs is 3.8 cm (1.5 in.). The fra m es are made up of flat stock mounted on the ribs in the central fuselage area. The 0.317-cm (0.125-in.) thick plexiglass window has an aperture of approxim a tely 38.1 by 76.2 cm (15 by 30 in.). The major difference in the simulation was in the depth of frames, stringers, and ribs.
The research-airplane depth is 3.8] cm (1.5 in.) and the test-fuselage depth was 6.35 cm (2.5 in.).
Final aircraft-designed containment, as well as the wire mesh of the window and initiation-system hardware, was used. The explosive severance was initi- ated by dropping a weight to actuate the control handle through a cable / pulley system.
A plexiglass witness panel was mounted inboard of the internal containment system at a distance of 21.6 cm (8.5 in.) from the mid-waterline (center stringer) of the egress opening. Dynamic pressure was monitored by two trans- ducers mounted on the plexiglass panel in the proximity of the expected loca- tion of the pilot's head in the experimental airplane. One transducer was mounted 5.1 cm (2.0 in.) aft of the forward internal containment, and the other was 20.3 cm (8.0 in.) aft. High-speed cameras (400 and 4000 pps) filmed the system operation and dynamics from the front, side, and rear.
P R OCE D URE Th e d es c rip t i o n o f th e d eve lop men t o f th e eg ress s y s t e m can b e l o gi cally b r o ke n into six ph ase s : syst e m s el ect i o n / d e v e l op m e n t co n sid e ra tions ; m at eri a l s se l e ct i on / s yst e m qualificat ion; i n itiation- s y s t em d eve lop men t ; e x p lo s i ve se v er - a nc e an d co n tainme n t d e v e lo pm ent; f u l l-sca le , flat-pa ne l tests; a nd a i rc r aft f u se la ge moc k -up t es t.
S y s t e m Sele ctio n / Devel o pmen t Co ns i dera t i o ns A mechanical or pyrotechnic system could provide the required egress open- ing. T he selection and development of this emergency in-flight egress system were based on the following: 1. Minimizing structural impact to the aircraft 2. Minimizing pilo t effort and response time to actuate 3. Minimizing system weight 4. Maximizing egress opening area 5. Minimizing pilot egress interference 6. Providing jettisoning force of egress panel 7. Providing passive, low-maintenance system 8. Providing maximum system reliability Materials Selection / System Qualification The pyrotechnic materials and techniques used in this egress system were selected on the basis of previous aerospace design experience and previously qualified components and systems.
Initiation-System Development The initiation subsystem was designed and developed with safety considera- tions paramount. The selection and development of the initiation system were based on the following characteristics: I. Independent system, isolated from onboard systems 2. Manageable actuation force 3. Fully secure in flight 4. Additional safetying measures on the ground 5. Positive "stops" to assure that the actuation is complete 6. Accessibility and reliability The complete initiation subsystem actuator (no detonators) was mounted on a flat-plate breadboard to evaluate the actuation forces required to overcome internal static and kinetic friction. The 89-N (20-1bf) maximum resistive force of the lanyard-actuated detonator was applied to the cable, and the pull forces necessary to overcome friction were measured.
Explosive Severance and Containment Development The development of the explosive severance technology progressed through several phases: I. Size the flexible linear-shaped charge and determine severance perfor- mance margins under worst-case conditions (a double thickness of aluminum and increasing the thickness beyond the expected limits). Also, determine the cut- ting performance of the FLSC and booster-cup combination inside the manifolds.
Past experience indicated that any foreign material such as potting in the c hevr o n area o f t h e F LSC destr o ys t he c u t t i ng efficien c y. Tests were conducted on double-thickness plates (1.016 on 1.016 mm (0.040 on 0.040 in.)).
2. Develop an external containment system to contain the explosive products outside the fuselage and provide a jettisoning force to the severed panel. A demonstration of the development was made by using small-panel test fixtures.
3. Develop a method of severing the central stringer in the egress area.
Again, small-panel test fixtures were used to demonstrate the local-area performance.
4. Develop a method of containing the explosive prod u cts inside the fuse- lage, assuring a performance margin. By using the internal containment devel- opment fixture described in the Apparatus section, containment tests were conducted under worst-case conditions of: a. 150 percent of the required explosive load was used, 3.19 g / m (15 grains / ft) RDX (cyclotrimethylenetrinitramine) instead of 2.13 g / m (10 grains / ft) b. No explosive pressure venting c. Minimum volumes d. Filling the volume with closed-cell, flexible foam (used to preclude contamination of the containment volume) e. Proximity of the FLSC to the aircraft and containment structure Full-Scale, Flat-Panel Tests To develop a n understanding of system-level performance, tests were con- ducted on the full-scale, flat-panel test mock-ups described in the Apparatus section. Performance parameters to be evaluated were complete severance, neatness / uniformity of severed edges, effect on aircraft structure during severance, capability of the containment structure (particularly at the stringer) to stop explosive products internally, jettison velocities and dynamics of the severed panel, and capability of the window mesh to prevent internal entry of the plexiglass window fragments on panel jettison.
Aircraft Fuselage Mock-up Test To demonstrate the final system design, a full-scale aircraft fuselage mock-up test was conducted. This test included actuating the mechanical ini- tiation subsystem with both detonators installed, the final design of the con- tainment system (particularly at the stringers) with closed-cell foam to pre- vent volume contamination, and the 3.8-cm (].5-in.) protective wire mesh on the inside of the window to prevent the plexiglass from moving internally. Inter- nal explosive debris and pressure were monitored with a plexiglass panel across the entire egress area.
_ S _ T S Sy s t e m S ele ct i o n / Deve lo pmen t Co ns i dera t i o ns Two candida t e e g ress s yst ems could meet t h e cons i dera t i o ns outlined in the P rocedure sec t ion - mechanical and pyrotechnic.
The mechanical system approach would require a large structural modifica- t ion and design effort to incorporate a door frame and doo r . T his would be followed by a release system, such as pulling hinge pins and actuating the latches, which could require considerable pilo t effort and time. The released door may then require (according to flight conditions) manual and aerodynamic jet t ison. Finally, a considerable effort may be required to validate the structural design of the modified aircraft under the high-stress, spin-pullout conditions.
T he p y r ot e c hni c s yst em appr o a ch w o uld use a flex i ble l i near-shaped charge (FLSC) to sever the existing skin and structure, following proven principles, applications, and materials. A minimal aircraft modification coul d be expe c ted, tha t is, a t taching the explosive and containment to the existing s t ructure. A pyro t echnic system would require little effor t t o initiate and, as a completely independen t energy source, would produce a highly responsive severance and jettisoning capability. Since t he expected structural changes would be minimal, no new load paths or structural analysis would be expected or necessary. Based on these considerations, the pyrotechnic system was selected for development.
Mate r ials Selection / Sys t em Qualification T he fle x ible line a r-shaped c harge ( F LSC) has been applied to several aero- space systems, including the F- I ll (ref. 1 ) and the B-] aircraft, in which the cockpi t is severed from the fuselage. The FLSC materials, organi c -precipita t ed hexanitros t ilbene ( H NS II) in a silver sheath, were developed specifically for thermal and age stability (ref. 2); applications include the F-J4, F-16, and A H -]G (Cobra) a i rcraft. The booster tip materials, H NS I in a steel cup, are applied almost universally to aircraft explosive t ransfer systems. The lanyard-actuated detonator was qualified for the F-14, F- J 5, and the Space Shuttle Orbiter. The capability of the FLSC to withstand severe environments is demonstrated by its many applications. T he functional qualification was based on componen t and system developmen t , emphasizing performance margins described in subsequent sections.
Initiation-System Development T he initiation system developed in this effort is shown in figure 4. A 40 ° rotational stroke o f the hand l e assembly (9.9 cm (3.9 in.)) pr o duces 4.95 cm (1.95 in.) pulley rotation and cable withdrawals. T he cables thread t hrough guide tub e s to provide 90 ° redire c ted pulls on the lanyards of the de t ona t ors. Since a 1. 2 7-cm (0.5-in.) s t roke is required to actuate the de t onator, a margin of at least 3 to ] exists. Each cable is fi tt ed with a c levis f o rk to a d ap t to th e de to na to r and a ball which is captured by a plate on the pulley. The cable lengths were adjusted to prevent simultaneous engagement / actuation of the detonators, which would result in twice the load.
Safety features were incorporated in the system to prevent inadvertent system actuation on the ground and in flight. The handle and pulley assembly is secured to the mounting plate by a ball-release bayonet safety pin for ground safety. An aluminum shear pin, again securing the handle and pulley assembly to the mounting plate, prevents inadvertent actuation without the safety pin for flight. Furthermore, the left-side cover posts (fig. 4) act as motion stops; the upper-post stop prevents forward motion of the handle, and the lower post provides a stop to assure that full actuation has occurred. A cover plate protects the entire pulley assembly and cable / tube ends. The entire assembly is mounted on the forward frame beside the pilot, just aft of the instrument console. The handle is positioned just above the lower extremity and 2.54 cm (1.0 in.) aft of the instrument panel.
The initiation-system breadboard revealed that the 3.175-mm (0.125-in.)
diameter, pure-aluminum shear pin sheared at 137.9 N (31 ibf). Furthermore, the static friction of either cable (preloaded to 89 N (20 ibf)) required only 93 N (21 ibf) at the handle. Actual friction loads will be much less in the system, since an 89-N (20-1bf) load will occur only at maximum stroke of the detonator, which occurs dynamically.
Expl o si v e Se v eran c e and C o n t ainmen t D e v el o pmen t The 68 explosive tests are briefly outlined in table I to establish the explosive severance and containment approaches and performance margins. T he results of the explosive sizing and performance comparisons are shown in table If. In determining the ability to sever double-skin thicknesses, the FL SC, composed o f 2.125 g / m ( 1 0 grains / ft ) RDX ( cyc l otrimethy l enetrinitramine ) , could sever / break the 1.016- on ].60- n _ n (0.040- on 0.063-in.) aluminum, which provides a performance margin of 58 percent. Furthermore, a ]3]-percent margin is achieved by using an H NS II FLSC of 3. ] 88 g / m ( 1 5 grains / ft) instead of the 2 .125 g / m RDX, resulting in an overall performance margin approaching 1 00 percen t .
In evaluating the performance of the booster tip / FLSC combination under manifolds, it was determined that the increased quanti t y of explosive (due to the tips) easily ruptured a single skin thickness and, with mini m al potting, could rupture two skin thicknesses. An already initiated tear would progress through the relatively short lengths where FLSC penetration did not occur.
The effec t iveness of the external containment approach in the small-panel tests is shown in figures 5 and 6. The no-containment test (fig. 5) produced ragged edges (large deflections) on both the severed panel and the aircraft skin mock-up. The containment test (fig. 6) produced smooth, neat edges on both the panel and the aircraft skin. Although the containment-test panel weighed nearly three t imes as much, no loss in jettison velocity occurred com- pared with the no-containment test. T he external contai n ment is a ].60-mm (0.0 63- in.) th i c k col d -r o lled s t eel co v e r pla t e ( 3 . 81 cm (1.50 i n.) wid e , as required), separated from the skin by a 3 .15-mm ( 0.]25-in.) aluminum-plate standoff. T he bent-down por t ion closes the cavi t y and smoo t hs the surface, redu c ing aerodynami c drag. T he cavity between the steel and the skin was necessary to assure adequate defle c tion of a double-thickness skin to allow reliable severance / fracture. A ].60-mm (0.063-in.) cavity allowed only p a rtial severan c e. T he final design of the e x ternal con t ainment is shown in figure 7 on the aircraft fuselage mock-up of the egress system.
T o reliably sever the central stringer, two lengths of explosive were used. One length was laid along the skin, through a hole in the s t ringer (inside the bend radius of the channel), and across the leg of the c hannel; t he other length was laid around the stringer and matched into common booster t ips at each end. T his arrangemen t introduced two problems: de t ermining the reliabili t y of severing the double-thickness material (stringer and skin) with the larger standoff to avoid the stringer radius, and determining how to manage the s t ructural d a mage and deformation caused by the doubled quanti t y of explo- sive. Several t es t s with larger s t andoffs than require d indicated sufficient energy existed t o sever an d tear t he m aterial with an adequa t e margin. T he doubled quan t i t y of explosive required doubling the s t ru c tural atta c hment bol t s (2.54-cm ( ] .0-in.) centers) and using steel in t ernal con t ainment structure on the se v ered p a nel.
T he internal containment structure to protec t the pilo t from explosive p roducts is shown in figure 8. The cross-sectional lines indicate t he loca- tions of subsequent structural views.
The princi p les of the internal explosive containmen t are shown in figure 9 (se ct ion A-A). T he skin is severed by the explosive, causing the struc t ure to t he right to be je t tisoned downward with the panel. T he explosive products are c on t ained within the free volume formed by the stringer, the angle t o the r i ght, and the cover channel. A c o ver channel is use d to preven t a left rotation of t he c o ver plate and strin g er d u e to the explosive press u r e. T he two cover plates above the channel stiffen the channel and cover the gaps at t he cover- c hannel interfaces. T he reinforcemen t angle pre v ents shearing damage from the c lose proximity of the explosive. The closed-cell foam (95-percent air) pre- vents contamina t ion of the free volume. I f this volume were filled with water (no foam), considerable deformation of the containment s t ructure could occur, possibly causing pilot injury.
T he same a p proach as described above is shown in figure 10 (se c tion B-B), except a cover an g le attached to the aircraft ribs is used to prevent rota t ion rather than a channel. T he cover angle is curved to match the aircraft con- tour. T he severed / jettisoned portion of the structure is below the stringer.
Figure 11 (section C-C) and fig u re 8 show the compl i cated welded s t ainless- steel containmen t structure used a t the forward and af t sections o f the central s t ringer. T he s t ru c t u re a t taches to the frames above and below t he stringer.
I n order to maintain an in t ernal free volume to dissipa t e the explosive energy of the flexible linear-shaped charge aro u nd the stringer, the str u cture had to p roject inboard into the cockpit. Also, this structure had to accommo d a t e the pos t -assembly installation of the detonators and initiatlon-cable guide tubes.
An approa c h s i m i lar to se ctio n B - B i s s h own i n f i gure 12 (section D-D), except t hat the cover angle is attached to the frame. Furthermore, there is no need for a reinforcement angle; the explosive was not mo u nted against the frame.
Full-Scale, Flat-Panel T ests The full-s c ale, flat-panel tests confirmed a number of system principles as well as dete c ting system problem areas. T he explosively severe d edges of the skin and je t tisoned panel were smooth an d uniform. T he dynamics of the severe d panels were unifo r m an d pred i ctable. T he panel was smoothly release d and pit c hed horizon t ally; the base of the panel swung upward in the direction of motion. H oweve r , the external containmen t structure and skin detached from the panel in the areas beneath the doubled FLSC around the central stringer, allowing explosive gaseous products to enter the fuselage.
The gases, crea t ed on detonation of t he explosive, sharp l y load the entire area wi t hin the in t ernal explosive containmen t , causing t he skin to deflec t between the bol t at t achments to exhaust the gases. These gases are highly visible as flame and smoke. The flame is a secondary burning of the unreac t ed carbon on mixing with the air. T he flame duration in bo t h t ests was approxi- mately 2 1 mse c , an unlikely ignition source of even the mos t reactive materials.
The manifold a t tachments and the plexiglass window retention were inade- qua t e on the firs t test and were corrected on the se c ond test. The manifolds (at t ached only t hrough the skin) pulled loose; a t tachmen t s t hrough t he frame prevented detachment. The window broke up due t o panel / frame oilcanning , tossing several pie c es inboard. A 3.8- by 3.8-cm (].5- by ].5-in.) wire mesh was s t retched a c ross the window and at t ached to t he internal containmen t s t ruc- ture to elimina t e internal debris.
T he je tt ison capabilities demonstrated i n the flat- p anel tests (s m all and full scale) are summarized in table I II. Al t hough t he sys t em weigh t increased , the amount of energy delivered per uni t weigh t was consis t ent.
Aircraft Mock-up Test Figures 13 and 14 show the neat, predictable, severed edges of the skin and the frame around the opening crea t ed by t he internal con t ainment. No in t ernal debris was dete c ted by the witness panel or high-speed camera cover- age (4000 pps). T he internal pressures measured were 34.5 kN / m 2 (5 psi) and 1 7.7 kN / m 2 (2.5 psi) with a d u ration of less than 1.0 msec. These pressure levels c ompare favorably with measurements made on the Bri t ish aircraf t Jet Provos t Mk.5, which employs an explosive cord-ac t uated overhead- c anopy sever- ance sys t em. Measurements at the ches t level of dummies indicated pressures of 50.3 to 117 kN / m 2 (7.3 to 17 psi). H owever, pressure levels at the dummy ears (inside t he helmet) were approxima t ely 2 6.2 kN / m 2 (3.8 psi).
T he flame duration and jettison velocity were considerably improve d by t he room temperature-vulcanizing (RTV) compo u nd applicat i on on the external con t ainment. T he a c tual duration of the flame was less than 5 msec. T he I0 i m pro vem e n t o f th e containmen t prevented c arbon particle and air mixing and burning. Furthermore, the jettison velocity (table III) increased 28 percent due to the improved sealing of the explosive-gas pressure wave. The severed panel weighed 6.62 kg (14.6 ib) and achieved a velocity of 13.7 m / sec (45 ft / sec).
The severed panel was completely intact, except for the window, following the test (figs. ]5 and 16). The wire mesh prevented any window fragments from entering the fuselage. The frame created by the external containment was smooth and u niform, and the skin was securely attached in all areas.
C O N CLUSIO N S A pyrotechnic-actuated, in-flight egress opening has been developed and qualified for use in a light, general aviation research airplane. This system will allow the pilot to bail out from the left side of the airplane.
The egress system is simple and highly responsive, requiring minimal air- plane modifications to incorporate. A complete, full-scale aircraft fuselage mock-up demonstrated the ability of the system to create an opening of approx- imately 76 by 76 cm (30 by 30 in.), including the window, in the cabin side.
T he total system weight was ] ].]7 kg (24.6 ib). The opening was created by small-quantity explosives (flexible linear-shaped charge) which severed and jettisoned a 6.62-kg (14.6-ib) po r tion of the fuselage skin and structure at a velocity of 13.7 m / sec (45 ft / sec). The explosive products are contained, presenting no debris or sound / pressure hazard to the pilot. Furthermore, the opening created is neat and smooth, presenting a minimal interference poten t ial to the pilot on egress.
System reliab i lity has been demonstrated by previous aerospace system applications and by functional tests. The pyrotechnic components and perfor- mance principles have been qualified on aircraft systems such as the F-I]] and B-] escape modules. All f u nctional parameters have been tested for this appli- cation to demonstrate substantial performance margins (a greater capability than required to accomplish the desired function). T he system will require no maintenance, except for a 5-year replacement cycle on the detonator.
This egress system technology is applicable to any aircraft of similar construction.
Langley Research Center National Aeronautics and Space Administration H ampt o n, VA 23665 February 29, 1980 R EFE R ENCES 1. Sc hi m mel , M. L . : T he F - Ill Crew M o dule : Ma jo r Challenge for Thermally Stable Explosives. McDonnell paper presented at the Symposium on Thermally Stable Explosives at the Naval Ordnance Laboratory (White Oak, Maryland), June 23-25, 1970. (Available as MDC 70-019.)
2. Bement, Laurence J.: Application of Temperature-Resistant Explosives to NASA Missions. NASA paper presented at the Symposium on Thermally Stable Explosives at the Naval Ordnance Laboratory (White Oak, Maryland), June 23-25, 1970.
T AB L E I . - EXP LO SI VE D EVELOPMEN T TE STS Exp l o si v e siz i ng / severan c e t es t s ...................... 2 3 Full-scale mock-ups: Aircraf t ................................. ] T ABLE II .- EXPLOSI VE SEVERANCE TE S TS AND COM P AR I S O N ( a ) Ex p l o sive severance t es t s Material severed, mm (in.)
Explosive, g / m (grains / ft) 1.016 on 1.016 1.016 on 1.60 1.60 on 1.60 (0.040 on 0.040) (0.040 on 0.06]) (0.063 on 0.063) ].488 (7) RDX Yes No 2.125 (]0) RDX Yes Yes No (b) Explosive severance c o mparison (2024- T 4 tapere d plates) 2.125 g / m (10 grains / ft) RDX will cut 1.53 mm (0.0602 in.) _ 131-percent 3.188 g / m (15 grains / ft) H NS I I will cut 2.00 mm (0.0788 in.)_ increase TABLE III .- PANEL JETTIS ON C OMPARISO N S Weigh t o f Vel ocity o f Ene r g y p er uni t we i gh t Ty pe t es t severe d panel , severed panel, o f s e v ered panel kg (ib) m / sec (ft / sec) J / kg (ft-lbf / ib) Sma l l-scale 0.34 (0.76) 9.8 ( 32) 48.2 ( ]5.9) (no c o ntainmen t) Small-s c ale 1.0 0 (2. 2 ) 1 0 . 4 ( 3 4) 53 . 2 (17.9) ( wi t h co n t a i nmen t) First mock-up 3.40 (7.5) 10.7 (35) 58.8 (]9.0) (flat panel) S econd m oc k -u p 5 . 9 4 (] 3 .1) 10. 7 ( 35 ) 57 . 2 (] 8 .9) (flat panel) F uselage moc k - up 6 . 6 2 ( ] 4. 6 ) 1 3 .7 ( 45 ) 9 4.0 ( 3] . 4 ) 0.76 mmR Sheath S(0,03 in.)
in,) core __0, mm 76mmR "_ (0.03 in.)
90o _+3°j 5.08 mm -= (0.20 in,) = Figure I.- Cross section of silver-sheathed HNS II flexible linear-shaped charge.
ACTUATION DIRECTION RELEASE SEAR COMPRESSIONSPRING (I.27-CM (.5 IN.) STROKE) 70 N / CM (40 LB / IN,) FIRING PIN ASSEMBLY PERCUSSION PRIMER OUTPUT CUP Figure 2.- Cross section of lanyard-actuated detonator.
L-80-113 Figure 3.- Internal view of aircraft fuselage mock-up.
L-80-114 Figure 4.- Mechanical initiation system.
L-80-]]5 Figure 5.- Small-panel severance test with no external containment.
L-80 I]6 Figure 6.- Small-panel severance test with external containment.
]7 L-80-] ]7 Figure 7.- External view of fuselage mock-up of egress syste m .
L-80-]]8 Figure 8.- Internal view of fuselage mock-up of egress system.
(Section views indicate subsequent illustrations.)
]8 6061-T6 (, 063 IN.)
7,6cM _(3,0 IN.)
.- . . . . ".. : '.
' " 3 8 CM 5 IN . ) REINFORCEMENT ANGLE. :_ ,, ._.:. ".L":'.
,16-CM 6061-T6 .... _- . "', : "- L CLIP (.063 IN.) ";':;; ' " ' " _ ' ''}' ON EXPLOSIVE ' I SKIN _.____- ,... L--.315-CM 2024-T4 (.125 IN.)
CLOS E D-C E LL F O A M .16-CM COLD-ROLLED STE E L (,063 IN_) Figure 9.- A-A and inverted A-A cross section of internal containment.
RIB SKIN STRINGER 12.7CM REINFORCEMENT ANGLE (5.0 IN.)
•16-CM 6061-T6 (.063 INS CLOSED-CELL FOAM .".'.":'i. :i {:. i:.i.: " "i (.063'16"CMIN)6061-T6 .315-CM 2024-T4 ____- 3.8 CM ------------_ (.125INS (1.5IN) Figure ] 0.- B-B cross section of internal containment.
IN ITIATO R M AN I FOLD \ . , C _]_ A I NMEN T __ BOOSTE R T IPS I NT ER NA L SK I M SHA PED C HARGE STRINGER .16-CM COLD-ROLLED STEE L iI '_ FLEXI B LE LI N EA R (.0 6 3 IN . ) _I !: CLOSED-CELL FOAM _ Figur e 11.- C-C c ross section of internal containment.
, 16-CM 5061-T6 ( , 063 I N.)
7,6 CM F (3,0 IN, ) ;. " .: i . " ," ,' ., .. ! -'i" ; - ; i_.!. " .; . ,, -.
FRAME , . ' . . : ' : . '."- ' :'.-, . .. ; . . - . " - 3[8 CM " " : ..:- ' : :': . _ 5 IN.)
,..._'L..'. i'" SPRING STEEL CLIP "'" ". ON EXPLOSIVE • , ° .
SKI N _ __-.- ,315 - cM 2024- T 4 F O AM _ _ (,125 IN') CLOSED-CELL ",16-CM COLD-ROLLED STEEL (,063IN.)
Figure 12.- D-D and inv e rted D-D cross secti o n of internal c o ntainm e nt.
2O i!! ' ii L-80-] ]9 Figure ]3.- Post-test external view of fuselage mock-up of egress system.
L-80-]20 Figure 44.- Post-test internal view of fuselage mock-up of egress system.
L-80-] 2] Figure ]5.- External view of severed panel.
L-80-] 22 Figure ]6.- Internal view of severed panel.
1 . Report No. 2. Government Acc essio n No. 3. Re c ipient's C a talog No.
N ASA TM-8 0 23 5 4. Title a nd Subtitle 5. Re po rt Da t e EMERGENCY IN-FLIGHT EGRESS OPENING FOR GENE R AL April 1980 AVIA TION AI RCRAF T 6. P er f or m ing Organ i z a tionCode 7 . Author(s) 8. Per f ormingOr ga n i zation Re po rt No.
Laur e n ce J . B e m e n t L-I 361 5 1 0 . Wor k U n i t No .
9. P er f ormingOrganizationName and Address 505-4 1 -I 3 - 05 NASA L angley Research Cen t er H a m pto n, VA 2 3665 11. C ontract o r Grant No.
1 3 , Type o f Re po rt and Period Covered 12. Spon s oringAgency Name a n d Address Technical M e mo r and u m Na tio na l A e r o na u tic s and S pa c e Ad mi nis t ra tio n Wash i ng to n, DC 205 4 6 1 4 . Sp o n so r i n g Ag e nc y Code 15 . SupplementaryNotes 16 . A b stract In s u ppor t o f a st all / sp in r e s e a rch pr o gram a t t h e NA S A L angl ey R e s e ar c h C en t e r, a n emergency in-fl i ght egress sys t em is be i ng ins t alled in a ligh t general aviation airplane. T he a i rplane has no prov i sion for egress on the lef t side. A lef t -s i de egress opening would grea t ly enhance the pilo t 's ability for ba i lout, part i cularly in a righ t sp i n. T o avoid a major s t ructural redesign for a mechanical door, an add-on ]].2-kg (24.6-ib) pyro t echn i c-ac t ua t ed system was develope d t o create an opening in the exis ti ng structure. The a i rplane skin w i ll be explosively severed around t he side window, across a central s t ringer, and down to the floor, creating an open i ng of approxima t ely 76 by 76 cm (3 0 by 30 i n.). T he severed panel w i ll be je t tisoned at an initial velo c i t y of approxi ma t ely 1 3 .7 m / se c (45 f t / se c ). Sys t em development i n c luded a to t al of 68 explos i ve severan c e t es t s on al u m i n u m ma t erial us i ng small samples, small and full-scale flat-panel aircraf t stru c tural mo c k-ups, and an a c tual a i r c raft fuselage. These tes t s proved explosive s i zlng / severance margins, explosive initia t ion, explos i ve produc t containment, and system dyn a mics.
This technology i s applicable t o any a i r c raf t of sim i lar construction.
1 7 . Key Words (Sug g estedby Author(s)) 1 8 . Distribution Statement Emergency egress Unclassified - Unlimited General avia t ion air c raf t E xplosively ac t uated Add-on sys t em Subje ct Category 0 3 19 . S ecurity Classif . ( o f thisre po rt) 20. SecurityCla ss if . ( o f this pa ge ) 21. N o . of Pa ge s 2 2 . Price" Unclass i fie d U n c la ssi f i e d 22 $ 4.0 0 * Fo r s ale b y t he Na t iona l Technic al I n fo rma t i o n Ser vic e, Spr i ngf i eld. V ir ginia 22161 N AS A - Langl ey , 1 9 80 National Aeronautics and SPECIAL FOURTH CLASS MAIL Postage and Fees Paid ,,_,C" ace Administration BOOK National Aeronautics and Space Administration NASA- 451 Washington,D.C.
20546
Official Business Penalty for Private Us e, $300 POSTMASTER: Po s tal M anual) Do Not Return _ A I f Und e liverable (S e ction 158