WO2012065155A1

WO2012065155A1 - Blade off protection systems and methods

Info

Publication number: WO2012065155A1
Application number: PCT/US2011/060550
Authority: WO
Inventors: James Jackson Milham Henry
Original assignee: Wintec Protective Systems, Inc.
Priority date: 2010-11-12
Filing date: 2011-11-14
Publication date: 2012-05-18

Abstract

A blade off protection system provides a pass through panel and a stop panel. The pass through panel and the stop panel are multi-layered panels that provide one or more fire resistant layers, energy absorbing layers, and ballistic layers. The pass through panel utilizes adequate layers to absorb a portion of the energy from a separated blade, while still allowing the separated blade to pass through the pass through panel. In contrast, the stop panel utilized adequate layers to prevent the separated blade from penetrating through the stop panel. The blade off protection system may optionally include a metal foam to further assist in absorbing the impact energy from a separated blade.

Description

BLADE OFF PROTECTION SYSTEMS AND METHODS

RELATED APPLICATIONS

[0001] This application claims the benefit of U.S. Provisional Patent Application No. 61/413,251 to Henry et al., filed on Nov. 12, 2010 and U.S. Provisional Patent Application No. 61/437,407 to Henry et al., filed on Jan. 28, 2011, which is incorporated herein by reference.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH

[0002] Not Applicable

REFERENCE TO A SEQUENCE LISTING

[0003] Not Applicable.

FIELD OF THE INVENTION

[0004] This invention relates to jet engine turbine fan blade failure. More particularly, to improving the safety of aircrafts by capturing and containing fatigued, fractured or broken fan blades within the engine shroud when a fan blade comes off the turbine shaft during a blade off incident within a turbine engine.

BACKGROUND OF INVENTION

[0005] Jet engine turbine fan blade failure presents a potential threat to aircraft structural integrity as well as passenger and crew safety in modern jet powered aircraft.

[0006] Modern jet engines may include two major thrust producing components. The first is the turbine compressor that mixes air and fuel, which in turn produces the mechanical motive energy that causes the fan blade assembly to rotate and draw air through the engine to create mechanical thrust. The blades of the turbine compressor may turn at approximately 20,000 to 30,000 rpms.

[0007] The second thrust producing component is the fan assembly which large set of closely set blades. These large blades are often visible on modern passenger aircrafts. These large blades may be directly connected via a shaft to the turbine compressor, but the large blades turn much slower at approximately 2500 to 3000 rpms. The large blades generating thrust to move an airliner by drawing ambient air through the engine.

[0008] Blade off incidents occur when the turbine compressor blades or fan blades fail. The blade that has failed may cause severe structural damage to other components of the engine and aircraft, thereby compromising aircraft flight integrity. Government agencies (e.g. Federal Aviation Administration) and other safety organizations may require testing to certify the safety of jet engines. For example, the safety testing may require that the fragments resulting from a blade off incident do not penetrate the outer casing of the engine.

[0009] Blade off incidents may be caused by a number of factors such as, but not limited to, structural failure, fatigue, blade creep, unwanted objects entering the engine, and the like. Hair line fissures in the blades may result from fatigue or the like and cause structural failure. Blade creep is when a blade permanently deforms or moves under the influence of stress and may result in the blade contacting the casing of the engine and failing. Blade creep may result from engines running too hot or too fast. Blade off incidents may also be caused by unwanted objects, like birds or the like, entering the engine and cause blade failure or blade separation.

[0010] In order to prevent the catastrophic dangers of blade off incidents, it is desirable to provide blade containment post blade separation. Current fan blade containment solutions are not fully satisfactory due to the high cost, lengthy fabrication and installation time, fabrication and installation complexity, weight, strength, temperature limitations, and the like.

SUMMARY OF THE INVENTION

[0011] In one implementation, a safety panel for turbine engine includes an adhesive securing a pass through panel to an inner or outer shroud of a turbine engine and multiple layers of material. The multiple layers of material may include a first layer comprising at least one layer of fire resistant material; a second layer comprising at least one layer of energy absorbing material; and a third layer comprising at least one layer of ballistic material.

[0012] In another implementation, a blade off protection system for turbine engine includes an inner shroud and a pass through panel coupled to a surface of the inner shroud. The pass through panel includes an first adhesive securing the pass through panel to the inner shroud, a first pass through layer providing fire resistance, a second pass through layer providing energy absorption, and a third pass through layer providing ballistic resistance, wherein the third pass through layer reduces a velocity of a separated blade while allowing the separated blade to penetrate through the inner shroud. The system also includes an outer shroud coupled to the inner shroud, wherein the inner shroud and outer shroud define a cavity; and a stop panel coupled to a surface of the outer shroud. The stop panel includes a second adhesive securing the stop panel to the outer shroud, a first stop layer providing fire resistance, a second stop layer providing energy absorption, and a third stop layer providing ballistic resistance, wherein the third stop panel prevents the separated blade from penetrate through the outer shroud.

[0013] In yet another implementation, a method for manufacturing a safety panel for turbine engines includes coating a multi-layered panel with a flexible coating to form a coated multi-layer panel, wherein the multi-layered panel comprises at least one fire resistant layer, at least one energy absorbing layer, and at least one ballistic layer; attaching the coated multi-layer panel to an adhesive to form a strip of a safety panel. One or more strips of the safety panel may then be applied to an inner shroud or outer shroud of a turbine engine.

[0014] The foregoing has outlined rather broadly various features of the present disclosure in order that the detailed description that follows may be better understood. Additional features and advantages of the disclosure will be described hereinafter.

BRIEF DESCRIPTION OF THE DRAWINGS

For a more complete understanding of the present disclosure, and the advantages thereof, reference is now made to the following descriptions to be taken in conjunction with the accompanying drawings describing specific embodiments of the disclosure, wherein:

FIG. 1 is an illustrative implementation of a cross sectional view of a jet engine;

FIG. 2 is an illustrative implementation of a cavity between an inner shroud and outer shroud;

FIG. 3 is an illustrative implementation of a pass through panel;

FIG. 4 is an illustrative implementation of a stop panel;

FIG. 5 is another illustrative implementation of a blade off protection system for a turbine engine; and

FIG. 6 is another implementation of a stop panel.

DETAILED DESCRIPTION

[0015] Refer now to the drawings wherein depicted elements are not necessarily shown to scale and wherein like or similar elements are designated by the same reference numeral through the several views.

[0016] Referring to the drawings in general, it will be understood that the illustrations are for the purpose of describing particular implementations of the disclosure and are not intended to be limiting thereto. While most of the terms used herein will be recognizable to those of ordinary skill in the art, it should be understood that when not explicitly defined, terms should be interpreted as adopting a meaning presently accepted by those of ordinary skill in the art.

[0017] It is to be understood that both the foregoing general description and the following detailed description are exemplary and explanatory only, and are not restrictive of the invention, as claimed. In this application, the use of the singular includes the plural, the word "a" or "an" means "at least one", and the use of "or" means "and/or", unless specifically stated otherwise. Furthermore, the use of the term "including", as well as other forms, such as "includes" and "included", is not limiting. Also, terms such as "element" or "component" encompass both elements or components comprising one unit and elements or components that comprise more than one unit unless specifically stated otherwise.

[0018] Blade off incidents can cause catastrophic damage to a jet engine and the aircraft. The aviation industry is well aware of the problem of blade containment post blade separation. Numerous solutions have been proposed for fan blade containment. Some of these solutions utilize aircraft grade aluminum, steel alloys, and graphite epoxy reinforced with Kevlar in an outer containment shroud and/or between the outer and inner shroud. However, none of these solutions are fully satisfactory.

[0019] These solutions require extremely expensive molds to create these parts. They also require many highly skilled man hours to construct or fabricate and install properly. Further, the materials do not satisfactorily solve the problem because of weight, shear strength, tensile strength, temperature limits, and the like.

[0020] The blade off protection systems and methods discussed herein provide simplicity of design and application; a lightweight solution; temperature resistance and thermal durability and functionality; and effective blade containment.

[0021] FIG. 1 is an illustrative implementation of a cross sectional view of a turbine engine 10. Compressor blades 15 compress air, which may then be mixed with fuel and ignited. The combusted gases exit through turbines 20 and nozzle 25. Fan blades 30 may rotate relatively slow in comparison to compressor blades 15 and may be significantly larger than compressor blades 15. Fan blades 30 draw in ambient air and generate thrust by moving at least a portion of the air through and area around the compressor. Inner shroud 35 conforms very closely to the edge of fan blades 30. Outer shroud or nacelle 40 provides the outer shell of engine 10.

[0022] FIG. 2 is an illustrative implementation of a cavity between inner shroud 35 and outer shroud 40. A cavity 45 is provided between inner shroud 35 and outer shroud 40. Cavity 45 between inner shroud 35 and outer shroud 40 is designed to serve several functions. One of the functions is to provide an open space wherein a separated blade, in the event of a blade off scenario, would be allowed to pierce the inner shroud 35 and retained within cavity 45. A major reason that such functionality is desired is to allow the separated blade to "get out of the way" of the attached compressor blades 15, fan blades 30, and/or other components of engine 10. This is critical because having the separated blade in the pathway compressor blades 15, fan blades 30, and/or other components of engine 10 may cause damage to other blades, thereby resulting in additional blades separating or damage to other components of the engine.

[0023] Pass through panel 100 may be provided on at least a portion of the interior surface of inner shroud 35, and stop panel 150 may be provided on at least a portion of the interior surface of outer shroud 40. While pass through panel 100 and stop panel 150 are respectfully provided on the interior surface of inner shroud 35 and outer shroud 40, in other implementations, they may be provided on the exterior surfaces. In other implementations, pass through panel 100 and stop panel 150 may cover anywhere from a limited portion of an inner or outer shroud to an entire portion of the inner or outer shroud.

[0024] Pass through panel 100 allows a separated blade to pass through inner shroud 35 and pass through panel 100. However, pass through panel 100 also absorbs some of the impact from the separated blade and slows the separated blade. In contrast, stop panel 150 is designed to prevent or minimize potentially catastrophic damage to the engine and other parts of an aircraft by preventing the separated blade from passing through outer shroud 40 and stop panel 150. In some implementation, a Kevlar wrap 50 may optionally be provided on outer shroud 40. Kevlar wrap 50 serves as an additional layer of material to prevent a separated blade from passing through outer shroud 40.

[0025] Cavity 45 may be filled with energy absorbing material(s), which may also be lightweight, for blade nesting space and structural stiffness. The cavity material(s) may be arranged into various energy absorbing structures. For example, cavity 45 may be filled with aluminum, composite aluminum, epoxy/carbon fiber filler composites, or other materials arranged into a "honeycombed" structure or the like. In another example, cavity 45 may be filled with a composite lattice material such as, but not limited to, an iso-truss matrixes, such as a carbon fiber iso-truss matrix. Further, the energy absorbing structure of the cavity material(s) may be filled with an expandable foam or the like to add significant structural integrity.

[0026] In some implementation, foamed aluminum, which can be machined easily and is extremely light and tough, may be utilized to fill cavity 45. The foamed aluminum can be filled with a filler material such as a soft pourable lightweight, low durometer elastomeric polymer that is an open or closed cell urethane, urea, foamed epoxy or the like. This combination would allow the material in cavity 45 to function better in absorb the separated fan or turbine blade and to hold the separated blade in suspension, thereby preventing the separated blade or debris from doing any further damage. The combination of materials in the engine 10 would help to prevent loose debris from becoming dislodged, flying about, or coming loose and doing further damage to other components of an aircraft. It would also provide a solid yet malleable fan blade containment area that would also function as a temporary storage area until the engine can be replaced. [0027] As an exemplary illustration, consider a blade 3 lb blade rotating at 3000 rpm. If blade separation occurs, the blade may generate from 14,000 to 21,000 foot pounds of impact energy moving at 700 to 1000 feet per second. Note that in some cases, blades may be as heavy as 10 lbs or greater and the blades may be rotating at higher speeds so significantly higher impact energy may result. As a result of these features, a separated blade can "get out of the way" of blades and other components of the engine, while preventing damage to other components of an aircraft by retaining the separated blade between inner shroud 35 and outer shroud 40.

[0028] FIG. 3 is an illustrative implementation of a pass through panel 100. Pass through panel 100 is designed to allow projectiles to "pass through" the panel. While pass through panel 100 may comprise materials to absorb a portion of the kinetic energy of such projectiles, pass through panel 100 does not prevent the projectiles from "passing through." Pass through panel 100 serves the two vital functions of allowing the blade to pass through inner shroud 35 and reducing the velocity of the blade significantly without actually stopping the blade. Pass through panel 100 may comprise multiple layers of material. Note that the layers discussed below may comprise one or more layers of materials. Further, it will be recognized by one of ordinary skill in the art that the order of the layers may vary in other implementations. For example, the order of the layers may vary in accordance with the placement of pass through panel 100 on the interior or outer surface of an inner shroud.

[0029] An adhesive layer 110 secures pass through panel 100 to inner shroud 35. For example, adhesive layer 110 may be a high strength adhesive tape, such as an acrylic foam tape (e.g. 3M VHB tape), epoxy, carbon epoxy, or any other suitable adhesive. First layer 115 may provide one or more layers of fire resistant materials such as, but not limited to, Nomex®, Nextel™, or other FAA certified fire resistant woven or unidirectional fabric like materials. The fire resistant material may help to prevent the spread of fire than may result from a blade off incident. Second layer 120 may provide one or more layers of energy absorbing material, such as, but not limited to, shear thickening material, dilatant material, D3o™, Sorbothane®, Ultra high molecular weight polyethylene (UHMWPE), or the like. The dilatant materials viscosity increases with shear strain, which allows the material to absorb energy from the impact of a blade. Additionally, the dilatant material may also provide vibration dampening depending on the dilatant material selected. Third layer 125 may provide a ballistic material, such as, but not limited to, Dyneema®, Tensilon, Kevlar®, para-aramid fibers, Zylon, iso-truss material or matrix, or other ballistic material. The ballistic material also absorbs impact energy, typically more than the dilatant material, but it should be noted that the ballistic material preferably does not prevent the blade or projectile from passing through. In some embodiments, it may be desirable to have materials with additional beneficial properties, such as shock and vibration attenuating properties, in addition to the ballistics, energy absorption, and/or fire resistance requirements. For example, materials such as, but not limited to, D3o Sorbothane materials or other shear thickening fluid materials, and the like may be selected for energy absorption properties, as well as desirable shock and vibration properties.

[0030] First, second, and third layers 115, 120, 125 may be sewn together using a high strength thread or the like. First, second, and third layers 115, 120, 125 may be covered by a spray on or roller coated outer coating 130 such as, but not limited to, any number of urethane, urea, or other elastomeric flexible coatings. The three layers 115, 120, 125 coated with an outer coating 130 may then be secured to inner shroud 35 by adhesive layer 110.

[0031] While the illustrative implementation discussed provides three layers, it will be recognized by one of ordinary skill in the art that number of layers may be varied as desired. For instance, each of the three layers may include multiple layers of material. In some implementations, a single material may provide multiple desired properties (i.e. fire resistance, energy absorption, and ballistic). As such, some implementations may require fewer than three layers.

[0032] FIG. 4 is an illustrative implementation of a stop panel 150. Stop panel 150 is similar in construction to pass through panel 100, but with a very important difference. Stop panel 150 serves to stop the blade completely, rather than allowing the blade to pass through, so as to contain and prevent the separated fan or turbine blade from doing any further damage to the structural integrity of the surrounding components of an aircraft. Pass through panel 100, materials in cavity 45, and stop panel 150 must safely and effectively stop a separated blade and arrest its outward and forward movement within the area between the inner shroud area and the outer shroud or nacelle.

[0033] While the following layers are illustrated as individual layers, the layers discussed below may comprise one or more layers of materials. Further, it will be recognized by one of ordinary skill in the art that the order of the layers may vary as discussed previously regarding the pass through panel.

[0034] An adhesive layer 155 secures stop panel 150 to outer shroud 40. For example, adhesive layer 155 may be a high strength adhesive tape, such as an acrylic foam tape (e.g. 3M™ VHB™ tape), epoxy, carbon epoxy, or any other suitable adhesive. First layer 160 may provide one or more layers of fire resistant materials such as, but not limited to, Nomex®, Nextel™, or other FAA certified fire resistant woven or unidirectional fabric like materials. The fire resistant material may help to prevent the spread of fire than may result from a blade off incident. Second layer 165 may provide one or more layers of energy absorbing material, such as, but not limited to, shear thickening material, dilatant material, D3o™, Sorbothane®, Ultra high molecular weight polyethylene (UHMWPE), or the like. The dilatant materials viscosity increases with shear strain, which allows the material to absorb energy from the impact of a blade. Additionally, the dilatant material may also provide vibration dampening depending on the dilatant material selected. Third layer 170 may provide a ballistic material, such as, but not limited to, Dyneema®, Tensilon, Kevlar®, para-aramid fibers, Zylon, iso-truss material or matrix, or other ballistic material. While stop panel 150 is similar to pass through panel 100, it should be noted that the purpose of stop panel 150 is different (i.e. stop panel 150 must stop a separated blade from passing through). As a result, second layer 165 and third layer 170 of stop panel 150 may be made up of many additional layers of dilatant and/or ballistic materials in order to completely stop and or arrest the forward movement of the separated fan or turbine blade. Similar to the pass through panel, one or more of the three layers of material 160, 165, 170 may also be selected to provide additional properties that are desired in addition to fire resistance, energy absorption, and ballistic properties.

[0035] To further improve arresting or stopping abilities, the internal structure of stop panel 150 may optionally contain one or more flexible high strength structural elements 180 such as, but not limited to, ultra-high strength metallic woven wires or non woven metallic meshes in order to provide greater arresting or stopping force. For example, a wire based product such as, but not limited to, Hard Wire® manufactured by Hardwire LLC. Hard Wire® is lightweight, flexible, inexpensive, and has extremely high tensile strength, in addition to other desirable properties. The one or more structural elements 180 may be incorporated in various positions within stop panel, such between one or more of the layers 160, 165, 170 and/or above and below layers 160, 170. A conservative estimate of the redundancy factor for arresting or capturing a blade is 6X or more. As such, stop panel 150 can likely capture multiple blades in the same event.

[0036] First, second, and third layers 160, 165, 170 may be sewn together using a high thread or the like. First, second, and third layers 160, 165, 170 may be covered by a spray on or roller coated outer coating 175 such as, but not limited to, any number of urethane, urea, or other elastomeric flexible coatings. The three layers 160, 165, 170 coated with an outer coating 175 may then be secured to outer shroud 40 by adhesive layer 155.

[0037] FIG. 5 is another illustrative implementation of a blade off protection system for a turbine engine 200. In another implementation of a blade off protection system, a single panel may be utilized. Turbine engine 200 may provide an inner shroud 205 and fan blades 210. Note that for illustrative purposes, outer shroud is not shown. In contrast to the previous implementations, rather than using both a stop panel and a pass through panel, a single panel may be utilized.

[0038] In some turbine engines, a separated blade may be capable of passing through the rear of the engine with minimal damage to other components of the engine. Thus, it is preferable to prevent the separated blade from passing through the inner shroud 205 of turbine engine 200. At least a portion of inner shroud 205 may be covered by stop panel 215 to prevent a separated blade from penetration and/or deforming inner shroud 205. The separated blade is retained within inner shroud 205 and passes through the rear of turbine engine 200 with minimal damage to other components of the engine.

[0039] Stop panel 215 may be a panel as illustrated in FIG. 4. Additionally, FIG. 6 is another implementation of a stop panel. The stop panel is secured to inner shroud 310 and provides a multi-layered panel. In particular, stop panel may provide one or more adhesive layer(s) 320, fire resistant layer(s) 330, energy absorbing layer(s) 340, and ballistic layer(s) 350. One or more fire resistant layers 310. Adhesive layer 320 may be an acrylic foam tape (e.g. 3M™ VHB™ tape), epoxy, carbon epoxy, or any other suitable adhesive. Fire resistant layer 330 may be Nomex®, Nextel™, or other FAA certified fire resistant woven or unidirectional fabric like materials. Energy absorbing layer 340 may be a shear thickening material, dilatant material, D3o™, Sorbothane®, Ultra high molecular weight polyethylene (UHMWPE), or the like. Ballistic layer 350 may be Dyneema®, Tensilon, Kevlar®, para-aramid fibers, Zylon, iso-truss material or matrix, or other ballistic material. [0040] The pass through panel and stop panel discussed above greatly simplifies the manufacturing and assembly process. The pass through panel and stop panel would allow engine manufacturers to wrap or tape the panels onto an inner shroud or outer shroud. In some implementations, installation would be performed by hand using human installers. In other implementations, installation using a mechanized wrapping or banding machine may be performed. Machine application would likely ensure a consistent tension and accurate control during application. A region of the inner shroud and outer shroud needing coverage by pass through panels and stop panels would exceed the actual width of the turbine or fan blade. For example, in some implementations, the width of coverage from the panels may exceed the width of the blades by a factor 25% to 50 % on each side to ensure catching the errant blade or blades. In other implementations, the width of the pass through panels and stop panels could easily vary in width depending on any number of factors. For example, the width may vary with the sizes of different engines, the size of turbine blades, any other factors, or a combination thereof.

[0041] In some implementations, the pass through panel or the stop panel may be a single strip to allow the panel to be applied in multiple overlapping passes. In other implementations, the pass through panel or the stop panel may comprise multiple strips, thereby allowing the strips to accommodate the various contours of inner and outer shroud(s). In yet another implementation, the pass through panel or the stop panel may be a monolithic panel. For example, in some implementations, the panels may be applied in a helical manner or continuously wound with a gradual overlapping of the strip with each circumferential wrap. The strip may be wound onto the inner or outer shroud several times until a desired region of the shroud is covered. This application process is advantageous for various reasons, including easier and more controllable application than wider tapes; easier to control the needful overlaps, spacing and most importantly, tensioning or winding control during the application; and the multiple overlapping and layering effects makes it easy to control velocity reduction of the pass through panel as desired by simply wrapping fewer or more layers. Multiple thin layers of "tape wrap" have better impact and energy absorption properties due to the controllability factors inherent in the principle of controlled delamination as an energy absorber. Consider for example proven principles found in ballistic body armor, which indicate that a multiplicity of thinner layers absorb impact energy considerably better and more efficiently than fewer thick layers.

[0042] The blade off protection systems and methods discussed herein provide a unique and novel way to minimize the potential for catastrophic damage to an aircraft during blade off incidents. The blade of protection system is light weight, easy to install, self- adhesive or-self sticking, easily upgradable, and easily conformable to irregular surfaces.

[0043] Implementations described herein are included to demonstrate particular aspects of the present disclosure. It should be appreciated by those of skill in the art that the implementations described herein merely represent exemplary implementation of the disclosure. Those of ordinary skill in the art should, in light of the present disclosure, appreciate that many changes can be made in the specific implementations described and still obtain a like or similar result without departing from the spirit and scope of the present disclosure. From the foregoing description, one of ordinary skill in the art can easily ascertain the essential characteristics of this disclosure, and without departing from the spirit and scope thereof, can make various changes and modifications to adapt the disclosure to various usages and conditions. The implementations described hereinabove are meant to be illustrative only and should not be taken as limiting of the scope of the disclosure.

Claims

CLAIMS What is claimed is the following:

1. A safety panel for a turbine engine comprising:

an adhesive securing the pass through panel to an inner or outer shroud of a turbine engine;

a first layer comprising at least one layer of fire resistant material;

a second layer comprising at least one layer of energy absorbing material; and a third layer comprising at least one layer of ballistic material.

2. The safety panel of claim 1, wherein a number of layers of ballistic material provided in the third layer is sufficient to prevent a separated blade from penetrating the safety panel.

3. The safety panel of claim 1, wherein a number of layers of ballistic material provided in the third layer is insufficient to prevent a separated blade from penetrating the safety panel.

4. The safety panel of claim 1, wherein the adhesive is a high bond adhesive, very high bond (VHB) tape, ultra high bond (UHB) tape, acrylic foam tape, epoxy, or carbon epoxy.

5. The safety panel of claim 1, wherein the fire resistant material is Nomex, Nextel, or a FAA certified fire resistant material.

6. The safety panel of claim 1, wherein the energy absorbing material is D3o, Sorbothane, a shear thickening fluid, or iso-truss material or matrix.

7. The safety panel of claim 1, wherein the ballistic material is an array of para- aramid fibers, Kevlar, Dyneema, Tensilon, Zylon, or high molecular weight polyethylene (UHMWPE).

8. The safety panel of claim 1 further comprising an outer coating surrounding the first, second, and third layers, wherein the adhesive layer is secured to the outer coating.

9. The safety panel of claim 8, wherein the outer coating is a urethane, urea, or elastomeric material.

10. The safety panel of claim 1 further comprising at least one structural element, wherein the structural element comprises a flexible, high strength woven material.

11. A blade off protection system for a turbine engine, the system comprising:

an inner shroud;

a pass through panel coupled to a surface of the inner shroud, the pass through panel comprising, an first adhesive securing the pass through panel to the inner shroud, a first pass through layer providing fire resistance,

a second pass through layer providing energy absorption, and a third pass through layer providing ballistic resistance, wherein the third pass through layer reduces a velocity of a separated blade while allowing the separated blade to penetrate through the inner shroud;

an outer shroud coupled to the inner shroud, wherein the inner shroud and outer shroud define a cavity; and

a stop panel coupled to a surface of the outer shroud, the stop panel comprising, a second adhesive securing the stop panel to the outer shroud, a first stop layer providing fire resistance,

a second stop layer providing energy absorption, and

a third stop layer providing ballistic resistance, wherein the third stop panel prevents the separated blade from penetrating through the outer shroud.

12. The system of claim 11, wherein the first adhesive or the second adhesive is a high bond adhesive, very high bond (VHB) tape, ultra high bond (UHB) tape, acrylic foam tape, epoxy, or carbon epoxy.

13. The system of claim 11, wherein the first pass through layer or the first stop layer is Nomex, Nextel, or a FAA certified fire resistant material.

14. The system of claim 11, wherein the second pass through layer or the second stop layer is D3o, Sorbothane, a shear thickening fluid, or iso-truss material or matrix.

15. The system of claim 11, wherein the third pass through layer or the third stop layer is an array of para-aramid fibers, Kevlar, Dyneema, Tensilon, Zylon, or high molecular weight polyethylene (UHMWPE).

16. The system of claim 11, wherein the pass through panel further comprises,

a first outer coating surrounding the first, second, and third pass through layers, wherein the first adhesive secures the first outer coating to the pass through panel; and

the stop panel further comprises,

a second outer coating surrounding the first, second, and third stop layers, wherein the second adhesive secures the second outer coating to the stop panel.

17. The system of claim 16, wherein the first outer coating or the second outer coating is a urethane, urea, or elastomeric material.

18. The system of claim 11, wherein the stop panel further comprises at least one structural element, wherein the structural element comprises a flexible, high strength woven material.

19. The system of claim 11 further comprising:

a filler material disposed in at least a portion of the cavity defined by the inner shroud and the outer shroud, wherein the filler material is a metal foam.

20. The system of claim 19, wherein the metal foam is filled with a elastomeric polymer.

21. A method for manufacturing a safety panel for turbine engines, the method comprising:

coating a multi-layered panel with a flexible coating to form a coated multi-layer panel, wherein the multi-layered panel comprises at least one fire resistant layer, at least one energy absorbing layer, and at least one ballistic layer;

attaching the coated multi-layer panel to an adhesive to form a strip of a safety panel.

22. The method of claim 21 further comprising placing at least one structural element in the multi-layered panel, wherein the structural element comprises a flexible, high strength woven material.

23. The method of claim 21 further comprising sewing the fire resistant layer, the energy absorbing layer, and the ballistic layer together prior to coating the multi-layered panel.

24. The method of claim 21 further comprising adhering one or more strips of the safety panel to an inner shroud or an outer shroud of a turbine engine.

25. The method of claim 24, wherein the one or more strips are helically applied to the inner shroud or the outer shroud.

26. The method of claim 21, wherein a number of layers in the ballistic layer is sufficient to prevent a separated blade from penetrating the safety panel.

27. The method of claim 21, wherein a number of layers in the ballistic layer is insufficient to prevent a separated blade from penetrating the safety panel.

28. The method of claim 21, wherein the fire resistant layer is Nomex, Nextel, or a FAA certified fire resistant material.

29. The method of claim 21, wherein the energy absorbing layer is D3o, Sorbothane, a shear thickening fluid, or iso-truss material or matrix.

30. The method of claim 21, wherein the ballistic layer is an array of para-aramid fibers, Kevlar, Dyneema, Tensilon, Zylon, or high molecular weight polyethylene (UHMWPE).

31. A blade off protection system for a turbine engine, the system comprising:

a stop panel coupled to a surface of an inner shroud, the stop panel comprising, an adhesive securing the stop panel to the inner shroud,

a first layer comprising at least one layer of fire resistant material, a second layer comprising at least one layer of energy absorbing material, and

a third layer comprising at least one layer of ballistic material, wherein stop panel prevents a separated blade from penetrating through the inner shroud.

32. The system of claim 31, wherein the adhesive is a high bond adhesive, very high bond (VHB) tape, ultra high bond (UHB) tape, acrylic foam tape, epoxy, or carbon epoxy.

33. The system of claim 31, wherein the first layer is Nomex, Nextel, or a FAA certified fire resistant material.

34. The system of claim 31, wherein the second layer is D3o, Sorbothane, a shear thickening fluid, or iso-truss material or matrix.

35. The system of claim 31, wherein the third layer is an array of para-aramid fibers, Kevlar, Dyneema, Tensilon, Zylon, or high molecular weight polyethylene (UHMWPE).