WO2010132086A1

WO2010132086A1 - A device and a method of preventing and removing jet engine compressor ice build up by asymmetric thrust bleed air valve dithering

Info

Publication number: WO2010132086A1
Application number: PCT/US2010/001219
Authority: WO
Inventors: Fergus D. Smith; Fergus S. Smith
Original assignee: Smith Fergus D; Smith Fergus S
Priority date: 2009-05-12
Filing date: 2010-04-24
Publication date: 2010-11-18

Abstract

The present invention provides anti-icing system for a turbofan compressor to overcome icing conditions that occur within the core compressor stage of current gas turbine jet engine designs. The anti-icing systems generally promote the oscillation of the engine's core compressor environmental conditions, such as its temperature and/or its pressure gradient Doing so allows the engine to disrupt the melting, evaporation, and refreezing process. Thus, oscillating the temperature and/or the pressure gradient of the engine prevents ice from accreting within the turbofan's core compressor by inhibiting the cooling capacity of evaporation for a given location Additionally, by straining the shape of compressor components in anticipation of core icing, oscillating the throttle causes ice to be continually shed in a subcritical state. Finally, forbidden and asymmetrically dithered throttle settings may be utilized to prevent ice-accretion prone temperature gradients from exerting their maximum cooling effects.

Description

TITLE OF INVENTION:

A Device and a Method of Preventing and Removing Jet Engine Compressor Ice Build Up Bv Asymmetric Thrust or Bleed Air Valve

Dithering

Inventors: Fergus D. Smith, 275 Williams St., Brattleboro, VT 05301; Fergus S. Smith, 1 Bungalow Rd., South Londonderry, VT 05155.

Cross references to related applications:

[001] This PCT application is related to the U.S. filed provisional patent application having serial number 61/216,076 file on 5/12/2009 by the same inventors.

Reference to federally sponsored research or development: NA Reference to joint research agreements: NA Reference to sequence listing: NA

TECHNICAL FIELD

[002] The present invention relates generally to gas turbine jet engines, and, in particular, to turbofan jet engines, and, in greater particularity, to turbofan anti-icing systems.

BACKGROUND ART

[003] The need for a turbofan compressor core anti-icing system stems from numerous in-flight, uncommanded shutdown events have occurred with turbofan jet engines over the past twenty years. It has been determined by industry that many of these engine powerloss events have been caused by glaciated and/or mixed phase ice particle accumulation on static components inside of the compressor stage of the turbofan, such as but not limited to an engine's sensors, stators, shroud, compressor inlet, and/or its guide vanes. While supercooled liquid water has previously been identified as an icing threat to the outermost components of a jet engine compressor, solid ice particles were previously thought to pass through the engine harmlessly.[l,2]

[004] This is no longer believed to be the case, and the source of these solid ice particles appears to stem from fine particle outflow emanating from vigorous convective cloud activity, as shown in FIG. 9, particularly in tropical locations. These events have occurred at altitudes of greater than 22,000 feet, which has been determined by industry to be the absolute limit at which supercooled liquid water is known to exist. Thus any water particle activity above this elevation will be the result of completely and/or partially frozen ice particles. [1]

[005] Thus a distinct and new form of icing from frozen ice particles, as opposed to icing caused by supercooled water droplets, has been identified. A supercooled liquid droplet is a form of water that remains liquid well below the melting point of water, which is 0 ⁰C. This is due to the fact that water requires a nucleating agent to allow it to freeze at 0 ⁰C. Without a nucleating agent to serve as a seed crystal, water may reach temperatures well below 0 °C . The practical effect of this is that supercoooled water droplets may then suddenly freeze the moment that such an agent is made available, such as at the surface of an engine component.

[006] Supercooled droplets typically require stable atmospheric conditions to maintain their liquid state. This is because a volatile atmospheric environment provides the mixing that allows nucleating agents to interact with the supercooled droplets. This type of interaction accelerates the formation of solid ice particles. The stable conditions necessary for maintaining a preponderance of supercoooled liquid water droplets does not generally exist in turbulent thunderstorms and tropical depressions that are characteristic of frozen ice particle-based turbofan compressor core icing incidents.

[007] FIG. 10 is a diagram of the flight envelope of various jet engine shutdown events that appear to be caused by glaciated and/or mixed phase ice particles, as opposed to supercooled water. It is notable how the majority of these events occur in temperature conditions that are approximately 0° - 20 °F above International Standard Atmosphere expected temperatures (ISA). This upward shift in temperature is a reflection of the unusual power of the local convective storm activity, within which these events are occurring. The intensity of this activity allows clouds to thrust ice particles to greater altitudes than would be considered normal, as in FIG. 9.[I]

[008] During the 199Cs numerous uncommanded commuter transport engine powerloss events have occurred in engines using high bypass ratio turbofans. These all "occurred under instrument meteorological conditions (IMC) in cloud, at thrust levels between 90 percent and 100 percent continuous cruise power and with precipitation (rain, ice or visible moisture) and light to moderate turbulence reported. The uncommanded thrust reduction manifested itself initially by a gradual decay in the fan rotational speed and a final stabilization of the engine at a sub-idle [non-functioning] operating condition." [1] This condition has been termed a "rollback".

[009] Fortunately, all engine powerloss events were restored in flight, once altitude was reduced to 10,000'. The cause of these rollbacks was determined by industry to be accreted ice on the compressor's stators, 23A, 23B, 23C, 23D, 23E, and 27A, 27B, 27C, 27D, 27E, 27F, 27G, and 27H, as indicated in FIG. 2. These powerloss events occurred in ice particle meteorological conditions with the presence of little or no supercooled liquid water. The result of this investigation led to successful engine modifications by applying heat to the stators. According to the Boeing Corporation's Jeanne Mason, uncommanded commuter aircraft shutdown events by ice particles have now ceased for this aircraft type. [1]

[010] Similar events have also occurred with wide-body, large transport jet aircraft, such as the Boeing 747. However, measures taken to remedy the problem for this class of aircraft have proven to be ineffective. Once believed to be a unique problem for commuter jet aircraft, the recognition of ice particle accretion in wide-body jet aircraft high bypass turbofans has demonstrated that ice particle accretion within a jet engine is not an isolated phenomenon. [1] Ice appears to be accreting and then shedding, causing dangerous surge/stall conditions within the engine, rollback events, and damage to the rotors by passing ice. Often, the dislodged ice is also extinguishing the combustor as the ice build-up enters that stage of the engine. This is referred to as a flameout. [1]

[011] According to Mason, in the early 1980's a multiple engine, large transport powerloss event occurred at 28,000' and -40 ⁵C, the theoretical coldest limit for supercooled liquid water. Investigators argued that the engine airflow rate in flight, even at descent power levels, was calculated to be high enough to prevent sufficient residence time for the heat transfer to take place between the ice particle and the engine component surfaces. Furthermore, given the high engine temperatures that quickly rise as air progresses through the engine, even during descent, there did not appear to be any mechanism to refreeze the liquid downstream, once it had melted. [1]

[012] In response to this event, engine manufacturer's raised the engine rotor speed during descent. This measure appeared to be effective. Having apparently solved this problem, no further action was taken by industry to investigate the matter of ice particle related powerloss, until the commuter engine problems occurred in the 1990's. While certain idiosyncratic engine designs were identified as being prone to core compressor stage icing, engine powerloss events have continued. Thus, since the 1980' s, as many as 100 powerloss events have been attributed to ice particle based ice accretion on stationary components of high bypass ratio gas turbine jet engines. [1]

[013] These events have occurred in engines from multiple manufacturers, using engines of varying design, and being used on at least nine different aircraft types, including wide body large transport aircraft, commuter jet aircraft, and executive jet aircraft, such as the Beechjet. Despite significant industry experience dealing with external icing on the wings of aircraft, as well as the adhesion of supercooled water on engine inlets and bypass fan 21 surfaces, the aerospace industry has experienced difficulties in modeling the conditions inside the turbulent environment of a high bypass jet engine. Data indicates that both older generation engines, as well as newer, electronically controlled, high bypass ratio engines are subject to ice particle powerloss. [1]

[014] From a sample of 46 powerloss events, Mason et al. have shown that these incidents have occurred in three phases of flight: climb (1 event), cruise (17 events), and descent (28 events). [1] The preponderance of events occurring during descent is attributed to several factors.

[015] First, ice accretion generally, but not always, occurs below 0 ⁰C. Thus ice accretion is more likely to occur at the low temperatures characteristic of high altitude, from which a plane initiates its descent. While water is known to freeze at levels of about 10 ⁰C in moderate conditions of low relative humidity, Mason et al.'s mention of this phenomenon does not acknowledge the centra I ity of evaporation's role in ice particle accretion. [1]

[016] Conversely, McVey et al. of GE Aviation do acknowledge that "evaporation can bring the surface to the freezing point." [9] While this is partially correct, it fails to clarify the crucial role of sublimation and the imbalance between the partial pressures of water at the surface of an engine component and the water vapor in the moving air of the engine system. This relationship is what gives rise to the particular form of evaporation exhibited by jet engines. By McVe/s own admission, "these processes are not well understood."⁹ We will comment more on this crucial phenomenon below.

[017] Second, when air density is at a minimum, ice particles constitute a greater proportion of total airflow through the engine, known to industry as the "scoop factor" effect. [1] This is because ice crystals possess more momentum than air molecules. Thus they resist being redirected around the engine as the engine passes through the air.

[018] Third, once an engine changes its throttle setting, strain on the rotors and/or the stators/exit guide vanes will be released. The resulting subtle change in shape of these engine components, due to changes in RPM and/or airflow angle, will serve as a basis to dislodge any accreted ice, which is inherently inelastic by nature. Ice's rigidity will place a strain on the adhesion bond between itself and the engine component, thus helping to delaminate the accreted ice.

[019] This problem is compounded by the fact that dislodged ice will have the greatest chance of extinguishing the combustor when the combustor is at a low-power, idle setting. Similarly, a potential phenomenon known as bridging may also exist. This theory asserts that ice may simply balance itself on a compressor component. Changing the throttle thus causes the balance of the ice to become disrupted. Such occurrences are most probable after an extended flight, during which time a critical level of ice may have developed. As the throttle is reduced, a potentially dangerous concentration of ice is inadvertently shed by the compressor, resulting in the aforementioned deleterious effects that are commonly caused by compressor core ice accretion.

[020] Finally, lower engine metal temperatures also exist when air flow is at a minimum, thus minimizing the heat available by the engine to compete with the heat sink created by the water particles, ice particles, and accreted ice.[l] This and the aforementioned factors together create a wider window of conditions that are favorable for ice accretion and subsequent ice dislodgment during descent.

Demystifying the Engine Icing Problem

[021] Strapp et al.³ have reported the microphysical conditions within a convective cloud system while on a test flight in the vicinity of Little Rock, Arkansas. During this test flight, using an unmodified engine, an engine rollback occurred. Of note during this study was the observation of a very low ice particle mean mass diameter (MMD) of approximately 40 μm. [1,3]

[022] While measurements of this nature are unlikely to be precise [1], the presence of small ice particles is important because they are proportionally more susceptible to thawing with decreasing diameter. Small particles possess a larger ratio of surface area to volume. Therefore, more heat may be transferred in a given period of time to the ice particles, thus increasing the probability of melting inside of the engine. Additionally, smaller droplets possess greater surface energy than larger droplets. This difference occurs because a smaller sphere will change its surface angle more severely for the same chord length as a larger particle, thus placing greater stress on the hydrogen bonds between water molecules.

[023] Consequently, small particles are fundamentally more susceptible to melting, while larger particles present a greater risk of impinging the surface of a stator by virtue of their greater inertia. [2] Note, however, that the greater inertia of the larger particle may result in a greater probability for contact with the rotor as well. This would serve as a basis for melting more of those larger ice particles. These competing factors may converge to combine a common source of melted water within a relatively wide range of particle diameters.

[024] There is only one temperature at which liquid water and ice can coexist, 0 °C, other than clusters of liquid water on the surface of the ice. The amount of liquid water on the surface of ice can be inferred from the partial pressure of water vapor at the surface of ice.

[025] Vapor pressure is the equilibrium pressure of a vapor in thermodynamic equilibrium with its condensed phase in a closed container. All liquids and solids have a tendency to evaporate into a gaseous form, and all gases have a tendency to condense back to their liquid or solid form.

[026] The equilibrium vapour pressure is an indication of a liquid's evaporation rate. It relates to the tendency of particles to escape from the liquid (or a solid). A substance with a high vapour pressure at normal temperatures is often referred to as volatile.

[027] The vapor pressure of any substance increases non-linearly with temperature according to the Clausius-Clapeyron relation. The atmospheric pressure boiling point of a liquid (also known as the normal boiling point) is the temperature at which the vapor pressure equals the ambient atmospheric pressure. With any incremental increase in that temperature, the vapor pressure becomes sufficient to overcome atmospheric pressure and lift the liquid to form bubbles inside the bulk of the substance. Bubble formation deeper in the liquid requires a higher pressure, and therefore higher temperature, because the fluid pressure increases above the atmospheric pressure as the depth increases.

[028] The vapor pressure of a single component in a mixture is called partial pressure. For example, air at sea level, saturated with humidity at 20 ⁰C has a partial vapor pressures [sic] of 24 mbar of water, and about 780 mbar of nitrogen, 210 mbar of oxygen and 9 mbar of argon. [4]

[029] Vapor pressure is the sum of all of the individual partial pressures within a given system, and can be described with the following equation:

V = P₁ + P₂ + ... + P_N

Where V = Vapor Pressure, and P = the partial pressure for a given chemical.

[030] The essential point is that when a state of disequilibrium exists between the vapor pressure at the surface of a condensed material, such as water, and the vapor pressure in the air, then there will be condensation and/or vaporization to bring the vapor pressure at the surface into equilibrium with the vapor pressure in the air. Moreover, even when the vapor pressure at the surface of an engine component is equal to the vapor pressure of the air flowing through the engine at that local region, each individual partial pressure in the air will trend towards equilibrium with its counterpart partial pressure that exists for that substance at the surface of the liquid and/or solid. This will occur even if the vapor pressure in the air and the vapor pressure at the surface are in a state of equilibrium.

[031] In the case of moist air flowing through an engine, any disequilibrium that exists between the partial pressures of water in the air and at the surface of the engine component will trend towards equilibrium, even if the vapor pressure of the system as a whole is in a state of equilibrium.

[032] In a mercury barometer, ambient air vapor pressure pushes a column of mercury about 780 mm up a tube which has been evacuated. Atmospheric pressure varies from less than 686 mm Hg during a hurricane to well over 780 mm Hg in a high pressure zone. At the altitudes at which people live, atmospheric pressure declines with altitude by very roughly 25.5 mm of mercury per 1000", with a decreasing rate with increasing elevation. At an altitude of 30,000', the air pressure of air is about 30% of that at sea level.

[033] Ice at 0 ⁰C is completely coated in liquid water and has a partial pressure identical to liquid water at 0 ⁰C, 4.579 mm Hg. As ice becomes cooler than 0 ⁰C, the proportion of liquid water to ice on the surface goes down. At -63 °C, the partial pressure water at the surface of ice is only 0.0053 mm Hg. At 100 °C water has a partial pressure of about 780 mm Hg. This is equal to and competitive with the vapor pressure of ambient air. Thus water boils if heat is supplied at the bottom of the vessel because its partial pressure is equal to or greater than that of the atmosphere. It is essentially pushing back against the weight of the air, while the raised water tumbles back down due to imbalances in the convection of the system.

[034] In a stable (non-convective) air mass, the partial pressure of water in the air will tend to go to equilibrium with the partial pressure of water on the surface of rain drops or ice crystals. At equilibrium, the probability of one molecule of water evaporating is equal to one molecule of water condensing. Importantly, whereas the relative humidity of air in a cold, high cloud will be high, the absolute humidity of air at cold, high altitudes will typically be very low.

[035] The latter is because air masses at high altitudes lack the vapor pressure to maintain solid or liquid water in equilibrium to the same extent as denser, lower altitude air masses. Thus the mass of water will be low as the excess water precipitates and/or evaporates until equilibrium is reached. Meanwhile, relative humidity merely describes the percentage of water as a proportion of what would be required to saturate the air at a given temperature. With higher ambient air temperatures, greater absolute humidity levels are possible before saturation (precipitation) occurs. This is because the greater temperature increases the partial pressure of water at the surface of the water or ice particle. This competitive interaction at the surface of the particle is what dictates whether condensation or evaporation occurs. [036] Therefore, low absolute humidity can mean very low relative humidity in regions inside the heated section of the compressor. The total amount of water available is not only low to begin with, but the warm air in the engine increases the air's saturation capacity by increasing the surface water's partial pressure, while lowering the relative humidity of the air. While it is true that a turbofan compressor increases the pressure of air as it passes through, its compression will not be perfectly adiabatic. This is because the rotors of the compressor will themselves promote warming of the air flow from friction. This will occur in addition to any warming that is achieved by compression alone, as predicted by Charles' Law. This effect will be most forcefully expressed in an inefficient engine.

[037] Additionally, because ice crystals and water will be put under substantial centrifugal forces, the relative humidity of a given cross-section of the compressor will be greatest at the outermost base of a stator and lowest at the statoKs tip. This leads to a substantial drop in relative humidity inside of the engine at certain areas. Cooling at these areas then can then be thermally conducted through the stator itself and serve as the basis to refreeze melted water that comes into contact with the stator. This water exists at higher concentrations near the base of the stator. Thus stratification of the water content within the compressor serves to increase in the airflow's capacity for evaporation and freezing as it passes through the warm engine, despite the high partial pressure of water vapor in the air in the vicinity of the base of the stator.

[038] Low relative humidity is why a person's throat typically feels dry while flying inside of a plane. The air is dry. This is also why ice accretion is not usually a problem in the stratosphere, above the tropopause, which usually serves as the upper limit of convection that occurs in the troposphere we live in. There typically is insufficient particulate water or ice to compete with the engine to begin with. However, once ice particles are thrust to unusually high altitudes, the higher partial pressure created by the warmer temperatures within the compressor, combined with the low vapor pressure of the atmosphere, will jointly promote evaporation and sublimation to a greater degree than that which occurs at lower altitude.

[039] This effect is one reason why planes are able to restart their engines after dropping in altitude. Because there is more water vapor in the air, the rate of evaporation is blunted. Thus the basis for ice accretion is impeded, and the engine is more able to compete with the ice crystal threat.

[040] During a violent thunderstorm, you can often see a flat top on the thunder cloud, thus it is said to have an "anvil" shape. That flat top is due to the tropopause, the dividing point where the stratosphere begins. Since temperatures generally cool down as you rise in the troposphere, but reverse, heating up as you rise in the stratosphere, convection generally ceases at the tropopause. Occasionally, extremely strong upward convection actually punches up through the tropopause and into the stratosphere, as shown in FIG. 9, because it has so much kinetic energy. This is why reported turbulence is relevant to understanding this problem. Its occurrence indicates unusually powerful convection and thus the availability of ice crystals at remarkable altitudes.

[041] At the top of the stratosphere (the stratopause) the temperature gradient reverses again with temperature falling as you go up in the ionosphere. The tropopause boundary is higher in warm weather, thus its ceiling varies widely between about 60,000' at the equator and about 30,000' at the Earth's poles. In the Northern United States it can vary from 30,000' to 50,000', depending on the temperature. This explains why engine powerloss events are occurring at higher than International Standard Atmospheric temperatures. By extending the tropopause to higher levels than is normal, planes are encountering instrument meteorological conditions (IMC) at higher altitudes than would normally occur otherwise.

The Role of Latent Heat

[042] So, an ice particle, starting at -40 °C for example, that adheres to an engine component surface after becoming liquefied to the melting point of water, 0 °C, will experience substantial evaporation, as well as sublimation, once the particle has stopped moving along with the airflow. Low local humidity at the surface of the water and/or ice particle would be maintained in such a state by the constant airflow, which would prevent a local partial pressure equilibrium from developing at the surface of the particle.

[043] Therefore, it must be understood that the latent heat of vaporization serves as a basis for a considerable potential energy differential between a water droplet and the surface that it adheres to. Once this liquefied and/or solid ice particle evaporates or sublimates, the subsequent thermal loss extracted from the engine component by evaporation and/or sublimation will be greater than the thermal input initially necessary to melt the ice. Thus, the net thermal transfer as a whole will be mass efficient and highly negative for any ice or water particle that adheres to the surface. While the temperature difference between a frozen ice particle and a warmer engine component surface partly contributes to the heat transfer caused by ice particles, the cooling effect caused by the transfer of latent heat during water's phase change dominates the freezing equilibrium in a wind-swept environment. The key is that a local equilibrium is unable to be maintained. Thus evaporation and sublimation continue unabated.

[044] To be clear, in reference to food, dieticians are referring to kiloca lories (1000 calories), which is often erroneously referred to as "calories" and written as "c", instead of "C". For the purposes of this discussion we are referring to individual calories as "c", and referring to degrees Celsius as "C". To change a gram of water 1 °C, the input or extraction 1 calorie of heat is required. This is called the specific heat of water. Thus, it is 1 calorie per gram per degree Celsius, or 1 c/g ⁰C. The specific heat of ice is 0.5 c/g ⁰C This is half the specific heat of water, and thus temperature changes of ice exchange only half as much heat as water itself.

[045] Water's specific heat is a high number, caused by hydrogen bonding between water molecules. Only ammonia has a higher specific heat. The specific heats of non-polar compounds tend to be considerably lower than that of water. Since heat is kinetic energy at the molecular level, and because polar molecules attract each other, this gives polar molecules an effective molecular weight that is higher than the real molecular weight. Because the actual weight in grams is in the denominator of the equation for specific heat, the specific heat of a polar liquid, such as water, is typically higher than for non-polar liquids. The reduced density and the reduced polarity of ice is why its specific heat is reduced, as compared to liquid water.

[046] Changing a substance's phase, such as turning it into a solid form, after beginning as a liquid form, requires a transfer of energy that is in addition to what is required to change the substance's temperature per its specific heat. To freeze a gram of water from 0 ⁰C liquid to 0 ⁰C frozen, 80 calories of heat must be removed, in other words, 80 c/g. This process is called the latent heat of fusion. The same amount, 80 calories, must be introduced to ice if you instead choose to thaw Ig of ice at 0⁰C into Ig of water at 0 °C.

[047] The heat involved is called "latent" heat because the temperature of water does not change while it is releasing or absorbing heat during its phase change. Phase change refers to when a substance changes its composition between solid, liquid, or gaseous states. To be clear, because the substance's temperature is not changing during this process, latent heat is not designated as per "C the way that specific heat is. A similar process occurs when 1 gram of water is vaporized, but the caloric transfer is about 600 c/g, or about 7.5 times greater than the latent heat of fusion.

[048] This is 600 times the specific heat of water, and 1,200 times the specific heat of ice ! This is a large number and a source of cooling that defies intuition in regards to icing in a jet engine. These numbers overwhelm the heat transfer contributed by the specific heat of ice itself, which is half that of water. The massiveness of this latent heat transfer greatly reduces the amount of water necessary to sufficiently cool a turbofan component, and it begins to explain what is missing from the aerospace industry's initial hypothesis about ice particle accretion.

[049] As described by Mason et al. (2006) in reference to empirical flight testing,

This flight test yielded the evidence that ice accretion may occur on engine component surfaces where local air temperatures under normal operation are significantly above freezing. The exact mechanism for icing accretion is not fully understood at this time, however, a hypothesis was developed based on the flight test results as follows. As soon as the engine enters glaciated/mixed phase conditions [completely frozen ice particles or partially frozen water particles], both liquid and ice particles coexist on the warm EGV [exit guide vane, which is a type of stationary engine component, and largely identical to a stator] surfaces. The presence of liquid on the surface slows down the ice particles, allowing heat transfer between the metal and ice particles to take place. [Emphasis mine] Heat removed from the metal reduces its temperature until the freezing point is reached, and ice forms. After this point, it is contended that further impingement of liquid and ice particles on the metal surface would accrete as ice even with local air temperatures higher than 0 degrees C. Accretion will continue to grow as long as both liquid water and ice particles continue to impinge on the ice surface, increasing the size of the blockage until the engine can no longer function properly. It should be noted that liquid water is a necessary condition locally at the EGV for ice accretion to continue. If there is no liquid water in the air stream, ice particles will bounce off the iced EGV surface in a way similar to what happens on the wing or the inlet of the engine. This phenomenon has also been observed in the mixed phase icing tests with supercooled liquid droplets and ice particles performed by Al Kahlil et al. (2003) in the Cox and Co. LeClerc icing wind tunnel. Since significant amounts of supercooled liquid water were not recorded by the weather instrumentation during the flight test, it is hypothesized that either only a very small amount of liquid water, ingested by the engine, is required for this process or that the liquid water impacting the EGV is actually produced by the melting of minute ice particles as they pass through the front section of the engine. [1]

[050] This is an incomplete theory, primarily because it does not account for the significant cooling effects caused by the latent heat of vaporization. While ice particles will contribute to seed the ice crystallization process on the surface of the engine component, the essence of Mason's et al.'s hypothesis proposes that frozen ice particles embedded in a thin film of water dominate the thermal transfer. However, as explained above, the specific heat exchange rate of ice is actually relatively feeble.

[051] The most common example of the role of the latent heat of vaporization occurs when a person sweats, thereby transferring heat from their body to the atmosphere. As described above, the caloric transfer during this process is dominated by the 600 calories required to vaporize a gram of liquid water, or 680 calories to sublimate a gram of ice. [052] The specific heat transfer of energy by water or ice is trivial in comparison, and the conductive effect of airflow on the surface of the water is also minor. The specific heat of air itself varies with the amount of water vapor, but this effect is most prominent when the air is saturated and contains water droplets, as in a cloud. The components of dry air, primarily nitrogen and oxygen, are non-polar, and at atmospheric pressure thus have a low specific heat, expressed as a function of mass, of only 0.24 c/g ⁰C.

[053] Additionally, the thermal conductivity for a material declines when it enters its gaseous phase. With less contact between molecules, less thermal transfer occurs. So air's thermal conductivity at 0 ⁰C is only about 0.024 W/mC (also expressed as (W * m) /( m² * 0 ⁰C)). Moreover, since air is a gas, thus low density, its specific heat, as expressed by volume, would be very low. The density of air at 1 atmosphere is only about 1.3 kg/m³ at 0 ⁰C, dropping even further with increases in altitude. With lower density there are fewer molecules to serve as a heat source.

[054] Also, it can be shown that the thermal transfer due solely by radiant heat emanating from the engine system will take an inordinately long time to melt even a small particle. This would be in the order of many hours or more. Therefore, this means that an ice particle will neither melt, nor freeze, just from travelling through the warm air of the engine. The poor thermal transfer of air to a small particle requires more time for melting or freezing to occur, and that time is not provided by a jet engine that swiftly processes vast quantities of air, and the particles passing through it, at high velocity.

[055] Conversely, once water or ice adheres to an engine component, forced convection friction promotes evaporation. While the friction may contribute the 80 calories or more of heat to melt a gram of ice at 0 °C, the partial pressure disequilibrium maintained by that same airflow would also encourage evaporation, thus extracting 600 calories per gram of evaporated water.

[056] Additionally, the inventors have hypothesized that ballistic effects are promoting the process of evaporation and/or sublimation. Because water is a polar molecule, water molecules are attracted to each other by virtue of the phenomenon known as hydrogen bonding. The result is that water prefers to aggregate into increasingly larger droplets because larger droplets cause less strain on the bonds between each molecule. While this effect can be readily observed during a rain storm on the front of a car windshield as small droplets coalesce into larger ones, this effect also occurs at the molecular level as well. The result is small puddles of water on a microscopic scale, and these puddles may be vaporized when oxygen and nitrogen molecules collide with these water puddles.

[057] This ballistic effect allows the liquid water to achieve a vaporized state more readily than it would so otherwise. This explains the observation by Mason et al. as to why there is freezing inside of the engine's compressor at inexplicably high temperatures. Rather than depressing the surface temperature of an engine component by a mere 10 °C,

Page U of 58 ballistic vaporization can conceivably reduce the surface temperature of water by an additional 20 ⁰C to 30 ^βC or more, instead.

[058] Ideally, any melted water would be expelled from a component's surface as water droplets. This would be preferable to the water evaporating as a vapor. However, ballistic effects and the existing partial pressure disequilibrium would cause the melted water to evaporate as soon as water puddles exist on the surface of the ice. Therefore, any melting that occurs from kinetic effects on the ice is overwhelmed by evaporative cooling. This effect is accentuated as the atmosphere's vapor pressure decreases with increasing altitude, reaching a maximum evaporation rate in a vacuum, such as outer space.

[059] This is an important factor for why dropping altitude is so beneficial for resolving these freezing events. With a lesser partial pressure disequilibrium, fewer particles evaporate. The result is that forced convection and warmer ambient airflow provide a greater and greater proportion of forced convection heat to the system, and less water evaporates or sublimates.

[060] This is also why these events continue to be possible, even at inexplicably high altitudes. With increasing altitude, the partial pressure disequilibrium only worsens, so that any ice particles that do manage to enter that environment are capable of maximum cooling activity. While an engine system is a volumetric system that will convolute this process, stratification of the relative humidity content in the air flow will provide the basis for the necessary low relative humidity conditions that promote evaporation.

[061] Therefore, the conductive heat transfer caused by air directly to an ice or water particle is negligible because of the above factors, and the heat transfer by means of latent heat is dominant. By way of example, if the relative humidity of air is below saturation when it rises, such as when it is pushed up the side of a mountain, it expands and cools at about 10 °C/1000m, or 10 °C/km. This is known as the dry adiabatic lapse rate, caused by the temperature decrease that occurs when a gas increases its volume.

[062] If the air is saturated, water vapor condenses because its partial pressure as a gas becomes greater than its partial pressure as a liquid at the surface of a droplet. This condensation delivers latent heat to the environment that it condenses to and/or within. Thus the air does not cool down as rapidly, and the wet adiabatic lapse rate is 6 °C/km, 4 °C/km lower than the dry adiabatic lapse rate. A drop in temperature with increasing altitude creates a partial pressure equilibrium that is in favor of condensation, and thus latent heat is delivered to the system in the process. This released heat minimizes the resultant temperature change.

[063] This process occurs in reverse as water evaporates. This is because water saturated air entering a jet engine will suddenly experience a steep drop in relative humidity within certain locations of the compressor. Ice particles typically would enter the engine in a state of equilibrium with the atmosphere. Raising the temperature of the environment alters this equilibrium in favor of evaporation and/or sublimation. This evaporation extracts heat from the surrounding environment, thus increasing the cooling capacity of the pre-existing water in, or introduced to, the engine system. This begins once liquid water or ice begins to adhere to the surface of an engine component. Otherwise, the water droplet or ice particle floats along with the airflow, relatively unaffected.

[064] Since ice always has some liquid water on its surface, that liquid can evaporate and be replaced by subtle, unobserved melting of the solid phase. Since we cannot easily observe the surface melting, but do observe the evaporation, it is called "sublimation". It is a misconception that melting does not occur during the process of sublimation. Since both melting and evaporation are involved, the latent heat of sublimation is simply 80 c/g + 600 c/g, or 680 c/g, a massive number. For an ice particle entering the engine system at -40 °C and then evaporating or sublimating, the heat transfer would be at least 80 c/g + 600 c/g + 0.5 c/g ⁰C * 40 ^βC, or 700 c/g, between the ice particle and a target surface. Of that 700 c/g, only 20 c/g, or 2.9% of that cooling would be due to the specific heat-related caloric transfer between of the ice particle and the warmer surface.

[065] This is why ice can accrete: By initially melting in a warm location, a water particle can then cool the substrate it comes into contact with by flowing and then cooling, due to partial evaporation, on another location, such as a stator. The shroud of the compressor would serve as an ideal conduit for this flow of highly humid air, given that the stators are connected directly to it, and given that the centrifugal force of the rotors would drive any liquid water from their blades directly to the shroud's surface. Unfortunately, some current computer simulation models oversimplify the ice accretion process by not including the shroud as a relevant engine design variable. [10]

[066] Thus, even though the rotor blades may possess a lower impingement rate for a smaller ice particle, the greater efficiency of the transfer of the resulting droplets by the shroud and to the stators would serve to compensate for the initially lower impingement rate by ice particles upon the rotors themselves. Once the surface of the stators or of the shroud is wetted, adhesion is made possible for additional ice particles. Also, the surface tension of an impinging water droplet decreases upon contact with the stator. This makes it easier for any airborne water droplet to freeze by increasing the surface area of the droplet on the engine component surface. Increased surface area promotes further evaporative cooling for a given volume of water.

[067] The relative ease with which smaller droplets remain melted is thus compensated for by the nucleating effects of the stator itself. This is particularly true when the stator is hydrophilic. This is because a hydrophilic surface will reduce the surface tension of the water even further, thus increasing the surface area that is available for evaporative cooling even more. And to reiterate, once the droplet adheres to the engine component, the airflow now acts upon and passes by the water, rather than passing along through the air. Thus forced convection and ballistic evaporation occurs, exerting the surface of the ice or water, as well as maintaining the partial pressure disequilibrium that promotes a maximum rate of evaporation.

[068] As such, an ice particle on the surface of a stator or of the housing will have a lower surface temperature than the previous rotor. This creates a temperature gradient reversal, which is the necessary precondition for turbofan compressor core ice accretion. This is because the stator experiences a wet bulb temperature depression by virtue of the effects of evaporative cooling. Passing airflow maintains a constant humidity level at the surface of the water or ice particle, thus preventing the water from establishing local partial pressure equilibrium between its liquid (or solid) phase and its gaseous phase.

[069] The wet bulb depression gradient will be at least between approximately 2 ⁰C and 10 °C at the surface of any water adhering to an engine component, and the previously known maximum depression of at least 10 °C, or possibly much higher from ballisticallγ induced vaporization, will occur locally at the greatest concentration of adherent water and/or accreted ice. The density of adherent water and/or accreted ice will directly correspond to the wet bulb temperature depression that is achieved.

[070] For example, an engine component that is 50% covered with ice will achieve a wet bulb depression of 50% of its maximum possible value. With a greater concentration of ice and/or water, there will be greater local evaporative cooling, and thus a maximum wet bulb depression in that local area. This wet bulb depression, which can be created by water without the benefit of ice, is why a constantly increasing ambient temperature gradient can be maintained by an engine, while nonetheless creating a temperature gradient reversal at the surface of the ice or water. This reversal is what allows ice accretion to happen.

[071] The competition between water vapor in the passing airflow with the partial pressure of water at the surface of the droplet or ice particle is what determines the partial pressure equilibrium. Passing this airflow quickly over the surface of the adherent water simply minimizes the change in concentration of water vapor in that adjacent air mass. This results in an otherwise constant water vapor partial pressure, and thus constant partial pressure disequilibrium. Rather than allowing water vapor to build up, and thus to deter additional evaporation, the newly evaporated water vapor is perpetually swept away.

[072] Thus, even if the engine's local air flow temperature is greater than freezing, the surface temperatures of any adhering water or ice will be lower than the local air flow temperature conditions surrounding the engine component. Because of the significant cooling effect caused by the latent heat of vaporization and/or sublimation, ice crystals per se are not an absolute requirement to create a heat sink at the stators. [073] Even under normal conditions, water itself can serve as a heat sink well before it drops to a freezing temperature. Once frozen, this ice then serves as the ideal nucleating agent for additional water particles, beginning the accretion process in earnest once the substrate is cooled to below 0 ⁰C. This increasing rate of ice accretion will occur at just below 0 °C substrate temperature. With lower temperatures creating lower partial pressures at the surface of the ice, less mass loss will occur from sublimation, resulting in an accelerating accretion rate. In the meantime, sublimation will still serve as a basis for evaporative cooling, while the latent heat of fusion will briefly release heat to the surface— and thus increase local temperatures briefly— as water turns to ice, thus releasing its latent heat of 80 c/g. So, temperatures will fluctuate around 0 ⁰C before experiencing a steady decline, once the water's phase change has been completed. This sequence is why the exit guide vanes' metal temperature represented in FIG. 37 follows the time series path that it does.

[074] At low enough relative humidity, an ice crystal can actually exceed 0 ⁰C and still remain frozen, as long as the caloric loss caused by the latent heat of vaporization sufficiently compensates for the higher temperatures. Instead of accumulating as melted water, there will instead be a loss of mass by the ice and/or water, which can then be replaced by additional impinging ice and/or water particles from the incoming air flow. If the mass flow of accreting water and ice particles is greater than the mass loss caused by sublimation, ice can then accrete on a surface that is warmer than 0 °C. Under normal conditions the upward temperature limit is approximately 10 ⁰C. However, due to the possibility of ballistic vaporization, substantially higher limits are theoretically possible, as described above. Once this process begins in earnest, the obstruction of the stators by accreted ice contributes to the overall effect by increasing the impingement rate of the ice particles.

[075] This is largely why in-flight empirical testing conducted during the 1990's documented a two stage freezing process of the engine's exit guide vanes (a kind of stator). [1 3] The water was cooling the surface by means of the latent heat of vaporization, and then nucleating the impinging liquid water particles, while being further nucleated itself by impinging ice particles. Once the wet bulb freezing temperature was achieved, this served as a catalyst for a greater ice and/or water particle adhesion rate, during which time accelerating engine deterioration was observed. This is because the rate of accretion of impinging particles became greater than the mass loss due to the sublimation of the accreted ice, which drops with decreasing temperature.

[076] Finally, there will also be a pressure decrease subsequent to the rotor, contributing to a measure of adiabatic cooling at the stator, as compared to temperatures at the rotor. So, if there is a reversal of the general warming trend through the engine, such as during the transition from a rotor to a stator, and if the rate of mass flow of particles deposits water or ice at a rate greater than the rate of sublimation, icing will accrete. If the engine's closed loop fuel delivery mechanism maintains a steady state, the ice accretion can eventually become critical at a specific location that is dictated by the temperature gradient of the engine. This is ultimately determined by the engine control mechanism itself.

[077] This is because the freezing window provided by the wet bulb depression will force ice accretion to occur in a similarly narrow section of the engine. By maintaining a steady state within the engine, so as to maintain constant thrust, the temperature gradient will remain steady as well. This encourages ice to occur in one narrow location. As ice accumulates, this wet bulb depression band will widen. Conversely, by changing altitude, the wet bulb region will shift away from the current location as the temperature gradient changes. This leads the airflow to process the accreted ice until it disappears from the location at which accretion originally occurred.

[078] Unlike the form of engine icing caused by supercooled liquid water droplets, dynamically manipulating the compressor's environmental conditions can control the accretion of ice caused by frozen ice particles. This is counterintuitive because supercooled liquid droplets can simply attach to any cold surface at the moment of contact. Changing the engine power will simply move the location of supercooled liquid water accretion, such as to a rotor, rather than eliminate it. This is because these particles are already below freezing, and thus they only require a nucleating agent, such as an engine component. This is because without a nucleating agent, even subzero water cannot readily establish the crystalline structure that is required for it to freeze.

[079] Ultimately, after the failure to eliminate ice particle powerloss events by setting the engines to run at higher temperatures during descent, or by attempting to divert ice to the bleed valve, it is clear that the aerospace industry has not determined how to manipulate the temperature gradient within the engine properly. Understandably, as industry experts have recognized, increasing the rotor speed within the engine will often only transfer the ice accretion point to a different location within the engine. [1] If this increased temperature is maintained in a steady state at precisely this shifted temperature, pressure, and airflow gradient, ice will simply accrete at an altered location but with the same result— uncommanded engine powerloss.

[080] Similarly, while compressor stage pressure increases can slightly lower the melting point of water by compressing and altering the crystalline structure of the ice, this will simply shift the ice accretion zone as well, rather than dictate whether ice accretion is made possible in the first place. Therefore, it is noteworthy that increasing pressure ratios between the low pressure compressor and high pressure compressor will not directly contribute to additional icing.

[081] Nonetheless, with increasing engine efficiency, a lower idle is capable of maintaining the necessary thrust during descent, relative to the size of the engine. This will result in a combustor that is relatively more vulnerable to being extinguished by any shed ice, by virtue of the lesser power requirement to maintain idle, and thus a relatively more vulnerable combustor flame.

[082] Conversely, because older engines have also experienced ice accretion, the culpability of the engine control system has been prematurely overlooked because of their superficial differences. The fundamental similarity of different engine control systems, be they mechanical, analog electronic, or digitally electronic, appears to have been underestimated. While the manner of the engine control differs, the result is generally the same: a stable maintenance of the temperature gradient of the compressor, the constancy of which allows for ice to accrete exactly in one location. This constancy maximizes the wet bulb temperature depression at that local environment. Ironically, the negative feedback control loop that modulates fuel flow, and thus maintains constant thrust from the engine, creates a positive feedback loop at the ice accretion zone within the compressor stage of the engine. This is especially true for any engine control methodology that lacks an intimate relationship with the airflow that passes through the engine.

SUMMARY OF THE INVENTION The Theoretical Basis for a Solution

[083] The solution to the ice particle accretion problem is to prevent the engine control system, in whatever its form, from maintaining excessively steady internal states within the compression stages of the engine. Despite the differing means by which various engine control systems attain their goal, the result is nonetheless identical: a steady engine state.

[084] A functional but nonetheless unsteady state can be achieved by dynamically oscillating the temperature, pressure, and/or airflow gradients within the engine and in concert with the plane's other engines. This will be more effective than permanently shifting the engine to a stable higher or lower temperature state.

[085] Like windshield wipers on an automobile, oscillating the temperature and pressure gradient within the engine continually drives the ice accretion zone away from previously accreted ice, into an area where the compressor previously was competitive with the effects of evaporation, only to then reverse itself and burn off what ice had been drawn to this previously immune area of the compressor. Additionally, by lowering the temperature over the rotors that are producing the liquid water, less water is available to bond or to accrete as ice in the original location where ice was originally accreting. This chokes off the bonding layer of water at its original source for the ice that is accreting for a given section of the compressor. Sublimation and forced convection thus are left to compete with the accreted ice that is left behind after the ice accretion zone has been shifted away. [086] In effect, the inventors are manipulating the fact that compressor core icing exists only as a relatively narrow band/cross-section of the compressor. Moving this band back and forth allows the ice immune regions adjacent to the ice accretion band, which is colder at one end of the compressor and warmer at the other end of the compressor, to compete with the accreted ice. This spreads a relatively constant volume of ice out over a larger region of the engine. While doing so will slightly alter the impingement rate of ice crystals to be higher or lower, depending on whether the throttle is decreased or increased respectively, the surface density and total thickness of accreted ice for a given surface area and time frame will be reduced.

[087] The greater the surface area of accreted ice for a given volume of ice that there is, the greater the rate by which the engine is able to compete with and eliminate the ice will be. The more ice that is induced into a normally immune area of the compressor, the lower the total ice accretion will be in a previously vulnerable region. The lower the density of ice on an engine component, the more compressor component surface area will be available for the compressor's airflow to conduct heat through the component and to the bonding layer of the accreted ice. The lower the density of ice, the greater the proportion of thermal infrared heat will be focused on the remaining ice. Because the rate of evaporation is directly and linearly proportional to the surface area of the evaporating material, the lower that the density of ice is for a given component's surface area, the lower that the average temperature drop will be for that total surface area will be from the effect of evaporation. This blunts the ability of evaporation to serve its critical role as the thermodynamic basis for compressor core ice accretion.

[088] Additionally, as the throttle is continually oscillated, the compressor components, such as but not limited to the stators/guide vanes and the rotors, will slightly alter their shape. This change in shape will cause ice to crack and to shed at subcritical levels because ice is an inelastic substance. Like shifting a balance beam under a gymnast, altering the throttle in anticipation of engine icing will allow the compressor component to destabilize the platform that is required for any accreted ice to maintain itself.

[089] Finally, the angle by which the airflow is delivered by the rotors will change when there is a change of throttle. This will serve as a basis to prevent the phenomenon known as bridging. Bridging is where ice becomes balanced on the leading edge of a stationary compressor component without actually becoming bonded to the component. This balance requires a stable airflow to allow this balancing act to be maintained. Oscillating the throttle prevents this balance from being maintained by causing the airflow to continually challenge the stability of the balance by buffeting the ice perpetually. Doing so thereby prevents bridged ice from ever reaching a critical state.

[090] The sum of the above stated effects leads to an exponential increase in the compressor's ability to compete with the ice. Therefore, oscillating the conditions within the compressor continually and deliberately in anticipation of possible compressor core ice accretion continually disrupts the ice accretion equilibrium, preventing its steady state from occurring as the ice accretion region is continually dithered from point to point. The lower density of ice serves to reduce the average wet bulb depression in a manner that is directly proportional to the surface density of the adherent water and/or accreted ice, thereby blunting the thermal transfer that is achieved by evaporation.

[091] Additionally, because in flight testing during ice particle meteorological conditions has determined that stationary engine parts are most subject to ice accretion, it can be inferred that moving the ice accretion zone away from stationary engine parts will cause these zones to cross over sections of the engine, such as compressor rotors, that are insensitive to ice particle ice accretion. The compressor's rotors, in particular, will be insensitive to ice accretion at these zones primarily because of the centrifugal force provided by their rotation. Thus, oscillating the temperature, pressure, and/or airflow gradient within the engine will allow the compressor rotors to serve as an additional means to disrupt the accretion of ice within the engine, while suppressing their contribution to the process by placing cooler air over the rotors and thereby limiting their ability to melt the ice particles that serve to bond other ice particles together.

[092] Accordingly, there is an established need for a turbofan compressor core anti-icing system that provides these features, in particular, the use of an oscillating throttle, an oscillating bleed air valve, and/or an oscillating engine environmental unit.

[093] The present invention is directed at a new anti-icing device and method for preventing and eliminating gas turbine jet engine compressor core ice accretion.

[094] The present invention further provides a novel device for oscillating the fuel flow, the bleed air valve to the bypass, the engine's load to the environmental control unit, the load from the electricity generation system, or any variable geometry, in such a manner that the engine's steady state can be altered.

[095] The present invention further provides control routines (150, 160, 170, 180) that may be used to alter the programming and/or the behavior of a plurality of engine control units (383), as well as a plurality of subroutines (161, 171, 181, 191) that are dependent upon the control routine's (150, 160, 170, 180) operation. These control routines (150, 160, 170, 180) may be added into or adjoined to an engine control unit (383), and may be used to include the necessary mathematical equations and/or software programming to govern the anti-icing oscillation system (110, 120, 130, 140, 382). This will require modifications to existing engine control systems, including but not limited to engine control units (383), autothrottle systems, and/or autopilot systems.

[096] The present invention further provides a novel device for adding precision to the anti-icing oscillation system. Subroutines are directed to control each engine in parallel to and in concert with the others, as shown in FIG. 15, such that the total thrust of all engines operating together is maintained constant, despite thrust oscillations by each individual engine.

[097] The present invention further provides a novel device to create for a uniform oscillation amplitude and frequency, or the oscillation be quantized such that the oscillation amplitude and/or frequency itself can be dithered randomly or deliberately, thus adding robustness to the anti-icing system.

[098] The present invention further provides that a predetermined temperature indicating the zone possessing the greatest wet bulb depression may be used as a basis to anticipate locations of compressor core ice accretion.

[099] The present invention further provides that by taking the absolute value of the difference between the high pressure compressor outlet temperature (HPCOT) and the low pressure compressor inlet temperature (LPCIT), or any two other temperature estimates from different locations within the compressor, the absolute value of the difference between the dry bulb equivalent to the wet bulb temperature at the zone within the compressor possessing the greatest capacity for core ice accretion (DBWBT) and the low pressure compressor inlet temperature (LPCIT), and creating a ratio of the latter divided by the former, the relative location of the freezing zone within the compressor can be calculated precisely. This ratio can be expressed with the following formula:

I DBWBT - LPCIT | / | HPCOT - LPCIT |

[100 The present invention further provides that by multiplying the above ratio, or another analogous formula, by the physical distance between the temperature input locations, the actual physical location of the freezing location can be inferred.

[101] The present invention further provides that zone of greatest ice accretion in the compressor, or the equivalent dry bulb temperature that exists in the same region of the engine, can be determined by referring to a database of empirically derived engine temperature and pressure relationships.

[102] The present invention further provides that the zone most favorable for stator and/or guide vane ice accretion can be deliberately forbidden or dynamically shifted to the compressor rotor blades, 24A-D and 28A-H. This can be achieved by altering the compressor operation settings, including but not limited to adjusting the fuel flow, adjusting the bleed valve to the bypass, adjusting the environmental control load, adjusting the electricity generation load, or adjusting any variable geometry inside or outside of the engine.

[103] The present invention further provides throttle and/or bleed air valve settings that result in zone of greatest compressor core ice accretion is made to skip over the stators and/or guide vanes, so as to avoid icing of the stators, 23A-E and 27A-H, or any other stationary engine components as well.

[104] The present invention further provides a basis for the use of forbidden throttle settings. This discourages ice accumulation by monitoring temperature gradients favorable to icing, and then deliberately avoiding them by manipulating the temperature gradients indirectly with the throttle, the bleed air valves, and or other means, such as but not limited to the environmental control unit (ECU).

[105] The present invention further provides that if the icing accretion zone cannot be directly estimated, alternating the anti-icing system's oscillation amplitude and frequency, based on a predetermined engine temperature gradient, can ensure that the most favorable icing accretion zone will exist on the compressor's rotors with at least equal probability to that of the stators.

[106] The present invention further provides a novel basis for adding robustness to the engine system by dithering of the steady state thrust levels of the engines asymmetrically in relation to each other. This means that each engine has a unique thrust demand placed upon it.

[107] The present invention further provides a novel electrically powered control system for analog dithering of a gas turbine jet engine throttle. This system utilizes an electrical oscillator HlA placed in line between the throttle IIOA and the engine system's fuel governor 112A. The preferred embodiment for the electrical oscillator HlA is an electronic timing circuit, such as a "one shot" circuit or an electronic timer. A switch 117 A, normally closed when the anti-icing oscillation system is not in operation, is used to control when the oscillator HlA is in operation. By intervening between the throttle IIOA and the fuel governor 112A, the fuel governor oscillator HlA can oscillate the thrust of the engine, thus oscillating the temperature and the pressure gradient within the compressor core.

[108] The present invention further provides a novel hydraulically powered control system for mechanical dithering of a gas turbine jet engine throttle. This system utilizes a hydraulic oscillator 13 IA placed in line between the throttle 130A and the engine system's fuel governor 132A. The preferred embodiment for the hydraulic oscillator 131A is a rotating hydraulic pump, such that the movement of the pump controls a valve which in turn controls the fuel governor.

[109] The present invention further provides a novel throttle-based anti-icing oscillation system. This system oscillates the throttle for (and fuel flow to) a plurality engines in concert with each other such that total thrust remains constant, while nonetheless eliminating compressor core ice accretion. [110] The present invention further provides a novel basis for an anti-icing oscillation system by oscillating the environmental control load on a plurality of engines. Once the environmental control load oscillation schedule is calculated 172, subroutines 173A, 173B, 174A, and 174B designated to each of one of the four engines are called to manipulate the environmental control load placed on each of these engines in concert with the other engines.

[Ill] The present invention further provides for the inclusion of an environmental control unit (ECU) on all of an aircraft's engines. This is necessary because not all engines possess an environmental control unit.

[112] The present invention further provides a novel basis for creating an anti-icing oscillation system by oscillating the throttle on a plurality engines in unison with each other, such that total thrust oscillates directly with the anti-icing system, while nonetheless eliminating compressor core ice accretion.

[113] The present invention further provides a novel basis for creating an anti-icing oscillation system by oscillating the bleed valve to the bypass on a plurality engines, each in concert with the others such that total thrust remains constant. This is achieved by allowing the behavior of bleed air valves to operate in a manner that is out of phase with respect to the bleed air valves of the other engines. The end result of the initialization process is an oscillation system that can manipulate the engine temperature gradient, while still maintaining constant thrust, by oscillating the settings to a plurality of bleed air valves.

[114] The present invention further provides a novel basis for creating a hybrid anti-icing oscillation system by oscillating the bleed air valve to the bypass, in conjunction with oscillating the fuel flow, on a plurality engines, each in concert with the others such that total thrust remains constant.

[115] The present invention further provides a novel basis for segmenting the throttle setting into two variables: a baseline throttle that is used in parallel by all engine systems, and an oscillating throttle, unique to each engine, that is controlled by the anti-icing oscillation system routines. This system allows for increases or decreases in engine throttle to meet changing flight requirements, while maintaining the ability of thrust- altering anti-icing systems to perform their function. A simple equation to clarify this concept follows:

Total Engine Throttle = Baseline Throttle + Oscillating Throttle

[116] It is commonly known to those familiar with the art of jet engine design that throttle may be controlled by a variety of means that includes but is not limited to basing an engine's thrust off of the compressor's RPM and/or the pressure ratio of the compressor inlet pressure and the turbine exit pressure. To adjust total throttle according to changing flight requirements, the baseline throttle can be adjusted such that the throttle of each engine is changed in unison, allowing the total thrust available to the plane to change in unison with it. If lower throttle is needed from all engines in equal amounts, the baseline throttle can be reduced, even as the oscillating throttle is operating independently of the baseline throttle. If greater throttle is needed from all engines in equal amounts, the baseline throttle can be increased, even as the oscillating throttle is operating independently of the baseline throttle.

[117] In operation, the present invention is a turbofan jet engine compressor core anti- icing system.

[118] An object of the present invention is to provide a basis to prevent and to eliminate ice accretion from within the compressor core of a turbofan jet engine.

[119] It is another object of the present invention to disrupt the steady state environment that exists within the turbofan's compressor core.

[120] It is a further object of the present invention to oscillate the temperature and/or the pressure gradient of the turbofan's compressor core.

[121] It is still a further object of the present invention to blunt the wet bulb depression that occurs in a turbofan engine when water and/or ice begin to adhere to a turbofan compressor core component.

[122] It is yet a further object of the present invention to provide an unsteady platform within a turbofan's compressor such that subcritical levels of ice are continually shed. This will occur as a result of the core compressor's rotors 24,28 and/or stators 23,27 altering their shape, due to changes in the strain caused by alterations in the throttle and/or bleed schedules of the engine 50, 60, 384. Because the shape of a rotor 24,28 and/or a stator 23,27 depends on the RPM and/or the airflow through the engine 50, 60, 384, fluctuating the engine's 50, 60, 384 environmental state will have the secondary effect of causing the geometry of the rotors 24,28 and the stators 23,27 to be slightly but continually altered. This will cause subcritical levels of ice to be cracked and subsequently shed.

[123] It is yet a further object of the present invention to statically dither the settings of a system of turbofan engines such that each engine's steady state differs from that of each of the other engines.

[124] These and other objects, features, and advantages of the present invention will become more readily apparent from the attached drawings and the detailed description of the preferred embodiments, which follow. BRIEF DESCRIPTION OF THE DRAWINGS

[125] The preferred embodiments of the invention will hereinafter be described in conjunction with the appended drawings provided to illustrate and not to limit the invention, where like designations denote like elements, and in which:

[126] FIG. 1 is a simplified schematic for the compressor stage of a prior art conventional gas turbine jet engine; [1,5]

[127] FIG. 2 is a simplified schematic for the compressor stage of a prior art conventional gas turbine jet engine. [5] This provides additional detail to what was presented in FIG. 1;

[128] FIG. 3 is a simplified schematic for the compressor stage of a prior art gas turbine jet engine. [5] This provides additional detail to what was presented in FIG. 2. However, instead of presenting engine parts, it describes areas of the engine that are prone to particular forms of icing;

[129] FIG.4 is a representation of the airflow relationship between a prior art jet engine's rotors and its stators; [2]

[130] FIG. 5 is a frontal view of a typical wide-body Boeing 747 jet airliner; [7] [131] FIG. 6 is an overhead view of a typical wide-body Boeing 747 jet airliner; [7]

[132] FIG. 7 is a simplified schematic of a prior art air conditioning system for a jet airliner; [8]

[133] FIG.8 is a simplified schematic of a novel air conditioning system for a jet airliner;

[134] FIG. 9 is an illustration of the airflow in a convective cloud storm system as it relates to the flight of a jet airliner through such a meteorological storm system; [6]

[135] FIG. 10 is a diagram of the flight envelope of various jet engine shutdown events that appear to be caused by glaciated and/or mixed phase ice particles, as opposed to supercooled water; [1,5,6]

[136] FIG. 11 is a flow chart of a novel power control unit for a fuel governor for a gas turbine jet engine anti-icing system containing four jet engines;

[137] FIG. 12 is a flow chart of a novel power control unit for a fuel governor for a gas turbine jet engine anti-icing system containing two jet engines;

[138] FIG. 13 is a flow chart of a novel power control unit for a fuel governor for a gas turbine jet engine anti-icing system containing four jet engines;

[139] FIG. 14 is a flow chart of a novel power control unit for a fuel governor for a gas turbine jet engine anti-icing system containing two jet engines; [140] FIG. 15 is a flow chart for a novel gas turbine jet engine throttle-based anti-icing system for a four engine aircraft. The purpose of this module is to alleviate icing conditions within the compressor section of the jet engine by oscillating and/or dithering of each engine's throttle settings;

[141] FIG. 16 is a flow chart for a novel gas turbine jet engine throttle-based anti-icing system for a two engine aircraft. The purpose of this module is to alleviate icing conditions within the compressor section of the jet engine by oscillating and/or dithering of each engine's throttle settings, while maintaining a constant direction by compensating with the plane's rudder;

[142] FIG. 17 is a flow chart for a novel gas turbine jet engine environmental control unit based anti-icing system for a four engine aircraft possessing at least one environmental control pack connected to each of four individual engines. The purpose of this module is to alleviate icing conditions within the compressor section of each jet engine by oscillating and/or dithering of each engine's environmental control load bleed air settings;

[143] FIG. 18 is a flow chart for a novel gas turbine jet engine environmental control unit based anti-icing system for an aircraft possessing at least one environmental control pack connected to each of two individual engines. The purpose of this module is to alleviate icing conditions within the compressor section of each jet engine by oscillating and/or dithering of each engine's environmental control load bleed air settings;

[144] FIG. 19 is a flow chart for a novel subroutine that is utilized by the anti-icing systems presented in FIG. 17 and FIG. 18;

[145] FIG. 20 is a flow chart for a novel subroutine that is utilized by anti-icing systems presented in FIG. 17 and FIG. 18;

[146] FIG. 21 is a flow chart for a novel gas turbine jet engine throttle-based anti-icing system for a four engine aircraft. The purpose of this module is to alleviate icing conditions within the compressor section of the jet engine by oscillating each engine's total throttle settings in unison with the other, while maintaining constant elevation by adjusting the trim settings;

[147] FIG. 22 is a flow chart for a novel gas turbine jet engine throttle-based anti-icing system for a two engine aircraft. The purpose of this module is to alleviate icing conditions within the compressor section of the jet engine by oscillating each engine's total throttle settings in unison with the other, while maintaining constant elevation by adjusting the trim settings;

[148] FIG. 23 is a flow chart for a novel subroutine that is utilized by the anti-icing systems presented in FIG. 21 and FIG. 22; [149] FIG. 24 is a flow chart for a novel subroutine that is utilized by the anti-icing systems presented in FIG. 21 and FIG. 22;

[150] FIG. 25 is a flow chart for a novel gas turbine jet engine bleed air valve based anti- icing system for a four engine aircraft. The purpose of this module is to alleviate icing conditions within the compressor section of the jet engine by oscillating and/or dithering of each engine's bleed air valve to the engine bypass settings;

[151] FIG. 26 is a flow chart for a novel gas turbine jet engine bleed air valve based anti- icing system for a two engine aircraft. The purpose of this module is to alleviate icing conditions within the compressor section of the jet engine by oscillating and/or dithering of each engine's bleed air valve to the engine bypass settings, while maintaining constant direction by adjusting the plane's rudder;

[152] FIG. 27 is a flow chart for a novel subroutine that is utilized by the anti-icing systems presented in FIG. 25 and FIG.26;

[153] FIG. 28 is a flow chart for a novel subroutine that is utilized by the anti-icing systems presented in FIG. 25 and FIG.26;

[154] FIG. 29 is a flow chart for a novel subroutine that is utilized by the anti-icing systems presented in FIG. 25 and FIG. 26;

[155] FIG. 30 is a flow chart for a novel subroutine that is utilized by the anti-icing systems presented in FIG. 25 and FIG. 26;

[156] FIG. 31 is a flow chart for a novel subroutine that is utilized by the anti-icing systems presented in FIG. 25 and FIG.26;

[157] FIG. 32 is a flow chart for a novel subroutine that is utilized by the anti-icing systems presented in FIG. 25 and FIG.26;

[158] FIG. 33 is a flow chart for a novel gas turbine jet engine hybrid anti-icing system for a four engine aircraft. The purpose of this module is to alleviate icing conditions within the compressor section of the jet engine by oscillating and/or dithering each engine's bleed air valve to the engine bypass settings, as well as each engines thrust, to create a hybrid anti- icing oscillation system;

[159] FIG. 34 is a flow chart for a novel gas turbine jet engine hybrid anti-icing system for a four engine aircraft. The purpose of this module is to alleviate icing conditions within the compressor section of the jet engine by oscillating and/or dithering each engine's bleed air valve to the engine bypass settings, as well as each engines thrust, to create a hybrid anti- icing oscillation system, while maintaining constant direction by adjusting the plane's rudder; [160] FIG. 35 is a flow chart for a novel subroutine that is utilized by the anti-icing systems presented in FIG. 33 and FIG.34;

[160] FIG. 36 is a flow chart for a novel subroutine that is utilized by the anti-icing systems presented in FIG. 33 and FIG. 34;

[162] FIG. 37 is a diagram representing the temperature changes to the surface temperature of an exit guide vane during a test flight encounter with ice particle meteorological conditions. [1] This temperature time series is represented alongside the turbofan's low pressure compressor rotor speed, abbreviated as Nl; and

[163] Fig. 38 is a flow chart depicting how a switch located in the cockpit, and therefore available to the pilot, may serve as a basis for the pilot to turn the anti-icing system.

[164] Like reference numerals refer to like parts throughout the several views of the drawings.

DESCRIPTION OF EMBODIMENTS

[165] The present invention is directed at a gas turbine jet engine compressor core anti- icing system.

[166] Turning to the drawings, wherein like components are designated by like reference numerals throughout the various figures, attention is initially directed to Illustration Group 1, FIG. 1 which illustrates a spinner 10 according to the present invention.

Illustration Group 1

[167] FIG. 1 is a simplified schematic for the compressor stage of a conventional gas turbine jet engine. [1,4] The system is comprised of a spinner 10, a bypass fan 11, a plurality of core compressor stators 12, a plurality of rotors 13, a low pressure compressor stage 14A, a space 15 between the low and high compressors where bleed air valves are typically located, a high pressure compressor section 14B, the bypass 16A and 16B, and a high pressure compressor outlet 17. Section 18 indicates the section of a gas turbine jet engine that has been determined by industry to be prone to supercooled water particle ice accretion. Section 19 indicates the section of a gas turbine jet engine that has been determined by industry to be prone to glaciated and/or mixed phase water particle ice accretion.

[168] FIG. 2 is a simplified schematic for the compressor stage of a conventional gas turbine jet engine. [5] This provides additional detail to what was presented in FIG. 1. The system is comprised of a bypass fan 21; a bypass stator 22; a plurality of core low pressure compressor stators 23A, 23B, 23C, 23D, and 23E; a plurality of low pressure compressor rotors 24A, 24B, 24C, and 24D; a low pressure compressor bleed air duct 25A, low pressure compressor bleed air valve 26A, a plurality of core high pressure compressor stators 27A, 27B, 27C, 27D, 27E, 27F, 27G, and 27H; a plurality of highpressure compressor rotors 28A, 28B, 28C, 28D, 28E, 28F, 28G, and 28H; a low pressure compressor bleed air duct 25B; a low pressure compressor bleed air valve 26B; and a high pressure compressor outlet 29. The arrow 20 indicates the flow direction of supercooled, glaciated, and/or mixed phase water particles as they enter the jet engine inlet.

[169] FIG. 3 is a simplified schematic for the compressor stage of a conventional gas turbine jet engine. [5] This provides additional detail to what was presented in FIG. 2. However, instead of presenting engine parts, it describes areas of the engine that are prone to particular forms of icing. The arrow in section 30 indicates the entry of supercooled, glaciated, and/or mixed phase water particles. Section 31 indicates the jet engine section that industry has determined to be prone to supercooled liquid water droplet icing. This includes the inlet, the spinner, the bypass fan, the bypass stator, and the initial stages of the low pressure compressor 33, as described in FIG. 2. Section 32 indicates the section of the engine that industry has determined to be prone to glaciated and/or mixed phase water particle icing. This includes the low pressure compressor 33, as well as the early stages of the high pressure compressor 34.

[170] FIG. 4 is a representation of the airflow relationship between a conventional jet engine's rotors and its stators. [2] This representation is comprised of rotors 40, stators 41, arrows indicating the direction of inlet air flow 42, an arrow indicating the angular rotation 43 of the rotors 40 as they cut through the inlet air 42, arrows indicating the altered direction of the airflow 44 after it has been acted on by the rotors 40, and additional arrows indicate the corrected direction of the airflow 45 after its orientation has been restored by the stators 41. Take notice of the pointed direction of the stators 41 with respect to the airflow 44 from the rotors 40. The prominence of the leading edge of jet engine stators 41 with respect to the incoming airflow 44 represents a narrow icing gathering point. This, in combination with a positive feedback loop from the engine control system, can create a runaway ice accretion point on the leading edge of the stators 41, despite the small size of the stators' 41 leading edge.

Illustration Group 2

[171] FIG. 5 is a frontal view of a typical wide-body Boeing 747 jet airliner. [7] This representation is comprised of a starboard outermost engine 50, a starboard innermost engine 51, a port innermost engine 52, a port outermost engine 53, an elevator 54, and a rudder 55. The up arrows, 56A and 56B, indicate how the trim and/or the elevator compensates for when the engine system's total net thrust decreases. The down arrows, 57A and 57B, indicate how the trim and/or direct control of the elevator compensates for when of the engine system's total net thrust increases. The side arrows, 58A and 58B, indicate how the rudder 55 compensates for the yaw that occurs when only two engines are oscillating their thrust. When the starboard engine or engines, 50 and/or 51, have more thrust than the port engine or engines, 52 and/or 53, this will cause the plane to yaw to the port side. Positioning the rudder 55 to starboard will correct the yaw by directing the plane's tail to port 58A. Conversely, when the port engine or engines, 52 and/or 53, have more thrust than the starboard engine or engines, 50 and/or 51, this will cause the plane to yaw to the starboard side. Positioning the rudder 55 to port will correct the yaw by directing the plane's tail to starboard 58B.

[172] FIG. 6 is an overhead view of a typical wide-body Boeing 747 jet airliner. [6] This representation is comprised of a starboard outermost engine 60, a starboard innermost engine 61, a port innermost engine 62, a port outermost engine 63, an elevator 64, a rudder 65, and ailerons and flaps 69A and 69B. The side arrows, 68A and 68B, indicate how the rudder 65 compensates for the yaw that occurs when and if engine throttle oscillation causes one side of the jet to display more thrust than the other side. This state can occur when a plane oscillates only one engine on each wing, or when both engines on the wing oscillate in an identical manner. When the starboard engine or engines, 60 and/or 61, have more thrust than the port engine or engines, 62 and/or 63, this will cause the plane to yaw to the port side because of the greater acceleration of the starboard side, as represented by 66B and 67A. Positioning the rudder 65 to starboard will correct the yaw by directing the plane's tail to port 68A. Conversely, when the port engine or engines, 62 and/or 63, have more thrust than the starboard engine or engines, 60 and/or 61, this will cause the plane to yaw to the starboard side, as represented by 66A and 67B. Positioning the rudder 65 to port will correct the yaw by directing the plane's tail to starboard 68B.

Illustration Group 3

[173] FIG. 7 is a simplified schematic of a conventional air conditioning system for a jet airliner. [6] This system is comprised of two jet engines 7OA and 7OB, an isolation valve 71 to control the flow of air between the jet engine bleed air sources from each engine, valves 72A and 72B controlling the bleed air flow to the air conditioning packs 73A and 73B, recirculation fans 74A and 74B, a conduit for ground preconditioned air 75, a mix manifold 76, a trim air system 77 to adjust the air temperature for each individual region of the plane, and outlets to the flight deck 78A, the forward cabin 78B, and the aft cabin 78C.

[174] FIG. 8 is a simplified schematic of an air conditioning system for a jet airliner. [6] This system is comprised of four jet engines 80A, 8OB, 8OC, and 8OD; isolation valves 81A, 81B, and 81C; to control the flow of air between the jet engine bleed air sources from each engine, valves 82A, 82B, 82C, and 82D controlling the bleed air flow to the air conditioning packs 83A, 83B, 83C, and 83D, recirculation fans 84A and 84B, a conduit for ground preconditioned air 85, a mix manifold 86, a trim air system 87 to adjust the air temperature for each individual region of the plane, and outlets to the flight deck 88A, the forward cabin 88B, and the aft cabin 88C.

Illustration Group 4

[175] FIG. 9 is an illustration of the airflow in a convective cloud storm system as it relates to the flight of a jet airliner through such a system. [6] This diagram shows how water particles are thrust high into the atmosphere, providing a means for their delivery into the inlet of a jet engine, even at cruising altitude.

Illustration Group 5

[176] FIG. 10 is a diagram of the flight envelope of various jet engine shutdown events that appear to be caused by glaciated and/or mixed phase ice particles, as opposed to supercooled water. [1,4,5] This diagram represents known icing envelopes Appendix C Continuous Maximum 100 and Appendix C Intermittent Maximum 101, linear regression lines representing International Standard Atmosphere expected temperatures (ISA) 102A, International Standard Atmosphere expected temperatures (ISA) plus 10° F 102B, and International Standard Atmosphere expected temperatures (ISA) plus 20° F 102C, and a plurality of engine flameout events 103 as they relate to International Standard Atmosphere temperatures.

Illustration Group 6

[177] FIG. 11 is a flow chart of a novel throttle control unit for a fuel governor for a gas turbine jet engine anti-icing system containing four jet engines. This system is comprised of an anti-icing oscillation control system for each engine HOA, HOB, HOC, and HOD; throttles HlA HlB, HlC, and HID; fuel governor oscillators 112 A 112B, 112C, and 112D; fuel governors 113 A, 113 B, 113 C, and 113D; fuel pumps 114 A, 114B, 114C, and 114D to deliver the required fuel to the combustors 115A, 115B, 115C, and 115D at the core of each jet engine; sensors 116A, 116B, 116C, and 116D to provide engine information to the governors 113 A, 113B, 113C, and 113D about engine performance; and electrical communication 119A, 119B, 119C, and 119D to connect the system components together, including switches 118A, 118B, 118C, and 118D to connect the fuel governor oscillators 112A 112B, 112C, and 112D to the fuel governors 113A, 113B, 113 C, and 113D. [178] The switches to the fuel governor oscillators 112A 112B, 112C, and 112D are normally open, and the switches 118 A, 118B, 118C, and 118D to the fuel governors 113 A, 113B, 113C, and 113D are normally closed. The opposite switch arrangement, by closing the switches 118 A, 118B, 118C, and 118D to the fuel governor oscillators 112 A 112B, 112C, and 112D, results when the anti-icing oscillation control system is in operation. The fuel flow 117A, 117B, 117C, and 117D is represented by the bold, hashed arrows running through the fuel pumps 114 A, 114B, 114C, and 114D, and the combustors 115 A, 115B, 115C, and 115D. The fuel governor oscillators 112A 112B, 112 C, and 112D can use a one shot electrical circuit or an electrical timer to control the oscillation process.

[179] FIG. 12 is a flow chart of a novel power control unit for a fuel governor for a gas turbine jet engine anti-icing system containing two jet engines. This system is comprised of an anti-icing oscillation control system 120A and 120B for each engine; throttles 12 IA and 12 IB; fuel governor oscillators 122A and 122B; fuel governors 123A and 123B; fuel pumps 124A and 124B to deliver the required fuel to the combustors 125A and 125B at the core of each jet engine; sensors 126A and 126B to provide engine information to the fuel governors 123A and 123B about engine performance; and electrical communication 129A and 129B to connect the system components together, including switches 128A and 128B to connect the fuel governor oscillators 122A and 122B to the fuel governors 123A and 123B.

[180] This arrangement allows the option of selecting an oscillation mode for the fuel governors 123A and 123B, or running them normally without the interruption of the oscillators 122A and 122B. The switches to the fuel governor oscillators 122A and 122B are normally open, and the switches to the fuel governors 123A and 123B are normally closed. The opposite switch arrangement occurs when the anti-icing oscillation system is in operation. The fuel flow 127A and 127B is represented by the bold, hashed arrows running through the fuel pumps 124A and 124B and the combustors, 125A and 125B. The fuel governor oscillators 122A and 122B can use a one shot electrical circuit or an electrical timer to control the oscillation process.

Illustration Group 7

[181] FIG. 13 is a flow chart of a novel throttle control system for a gas turbine jet engine anti-icing system containing four jet engines. This system is comprised of an anti-icing oscillation control system 130A, 130B, 130C, and 130D for each engine; throttles 131A, 131B, 131C, and 131D; fuel governor oscillators 132A 132B, 132C, and 132D; fuel governors 133A, 133B, 133C, and 133D; fuel pumps 134A, 134B, 134C, and 1340 to deliver the required fuel to the combustors 135A, 135B, 135C, and 135D at the core of each jet engine; sensors 136A, 136B, 136C, and 136D to provide information to the governors 133A, 133B, 133C, and 133D about engine performance; hydraulic valves, including but not limited to a ball valve or a solenoid valve, 138A, 138B, 138C, and 138D to connect the fuel governor oscillators 132A, 132B, 132C, and 132D to the fuel governors 133A, 133B, 133C, and 133D, and hydraulic communication 139A, 139B, 139C, and 139D to connect the system components together.

[182] The valve outlets to the fuel governor oscillators 132A, 132B, 132C, and 132D are normally closed, and the valve outlets to the fuel governors 133A, 133B, 133C, and 133D are normally open when the anti-icing oscillation system is OFF. The opposite valve arrangement turns the anti-icing oscillation system is ON.. The fuel flow 137A, 137B, 137C, and 137D is represented by the bold, hashed arrows running through the fuel pumps 134A, 134B, 134C, and 134D, and the combustors 135A, 135B, 135C, and 135D. The fuel governor oscillators 132A, 132B, 132C, and 132D can use a rotating hydraulic piston to serve as a basis for the oscillation. This can then be connected by linkage to a solenoid valve to control the fuel flow and/or the fuel governor behavior 133A, 133B, 133C, and 133D.

[183] FIG. 14 is a flow chart of a novel throttle control system for a gas turbine jet engine anti-icing system containing two jet engines. This system is comprised of an anti-icing oscillation control system 140A and 140B for each engine; throttles 141A and 141B; fuel governor oscillators 142A and 142B; fuel governors 143A and 143B; fuel pumps 144A and 144B, to deliver the required fuel to the combustors 145A and 145B at the core of each jet engine; sensors 146A and 146B to provide information to the governors 143A and 143B about engine performance; hydraulic valves, including but not limited to a ball valve or a solenoid valve, 148A and 148B to connect the fuel governor oscillators 142A and 142B to the fuel governors 143A and 143B, and hydraulic communication 149A and 149B to connect the system components together.

[184] The valve outlets to the fuel governor oscillators 142A and 142B are normally closed, and the valve outlets to the fuel governors 143A and 143B are normally open when the anti-icing oscillation system is OFF. The opposite valve arrangement turns the anti- icing oscillation system is ON. The fuel flow 147A and 147B is represented by the bold, hashed arrows running through the fuel pumps 144A and 144B, and the combustors 145A and 145B. The fuel governor oscillators 142A and 142B can use a rotating hydraulic piston to serve as a basis for the oscillation. This can then be connected by linkage to a solenoid valve to control the fuel flow and/or fuel governor behavior 143A and 143B.

Illustration Group 8

[185] FIG. 15 is a flow chart for a novel gas turbine jet engine throttle-based anti-icing system 110, 120, 130, 140 for a four engine aircraft. The purpose of this system is to alleviate icing conditions within the compressor section of the jet engine by oscillating and/or dithering of each engine's throttle settings. This system is comprised of an anti- icing control routine 150 that governs behavior of subroutines that are nested within the program; a power ON or OFF conditional 151 that determines whether the thrust oscillation anti-icing control routine 150 is set to ON. When the power is OFF or returns with a NO, then the routine reverts back to the anti-icing control routine 150 unchanged. When the power is OFF or returns with a YES, then the program proceeds to a thrust oscillation schedule routine 152. This routine determines the engine throttle oscillation schedule according to the compressor stage temperatures and/or pressure gradients for each of the jet aircraft's engines, in addition, call commands 153A and 153B are directed to an engine control subroutine, Thrust Oscillation Sub One", depicted in FIG. 23. Thrust Oscillation Sub One governs the thrust oscillation for any engine that the routine is dedicated to by the anti-icing control routine 150. Similarly, call commands 154A and 154B are directed to an engine control subroutine, Thrust Oscillation Sub Two", depicted in FIG. 24. Thrust Oscillation Sub Two governs the thrust oscillation for any two engines that the routine is dedicated to by the antHcfng control routine 150. Call commands 153A, 153B, 154A, and 154B all run in parallel to each other. As each subroutine operates and the thrust of each engine is oscillated up and down, the net total thrust produced by all engines is held constant. The output of the anti-icing control routine 150 and the call commands 153A, 153B, 154A, and 154B are in communication with at least one engine control unit 159, which possesses control over at least one throttle 111, 121, 131, 141 to at least one engine 50, 60.

[186] FIG. 16 is a flow chart for a novel gas turbine jet engine control unit throttle-based anti-icing system 110, 120, 130, 140 for a two engine aircraft. The purpose of this system is to alleviate icing conditions from within the compressor section of the jet engine by oscillating and/or dithering of each engine's throttle settings, while maintaining a constant direction by compensating with the plane's rudder. This system is comprised of an anti- icing control routine 160 that governs behavior of subroutines that are nested within the program; a power ON or OFF conditional 161 that determines whether the thrust oscillation anti-icing control routine 160 is set to ON. When the power is OFF, or returns with a NO, then the routine reverts back to the anti-icing control routine 160 unchanged. When the power is OFF or returns with a YEiS, then the program proceeds to a thrust oscillation schedule routine 162. This routine determines the engine throttle oscillation schedule according to the compressor stage temperatures and/or pressure gradients for each of the jet aircraft's engines. In addition, a call command 163 is directed by the anti- icing control routine 160 to an engine control subroutine. Thrust Oscillation Sub One", depicted in FIG. 23. Thrust Oscillation Sub One governs the thrust oscillation for any two engines that the routine is dedicated to by the anti-icing control routine 160. Similarly, a call command 164 is directed by the anti-icing control routine 160 to an engine control subroutine. Thrust Oscillation Sub Two⁰, depicted in RG. 24. Thrust Oscillation Sub Two governs the engine oscillation for any two engines that the routine is dedicated to by the anti-icing control routine 160. Finally, a yaw adjust routine 165 corrects the plane's yaw by using the aircraft's autopilot. The call commands 163 and 164, and the yaw control routine 165 all run in parallel to each other. As each subroutine operates and the thrust of each engine is oscillated up and down, the net total thrust produced by all engines is held constant, and the aircraft is able to maintain straight and level flight. The output of the anti-icing control routine 160 and the call commands 163, 164 the yaw adjust routine 165 are in communication with at least one engine control unit 169, which possesses control over at least one throttle 111, 121, 131, 141 to at least one engine 50, 60.

Illustration Group 9

[187] FIG. 17 is a flow chart for a novel gas turbine jet engine environmental control unit- based anti-icing system 110, 120, 130, 140 for a four engine aircraft. The purpose of this system is to alleviate icing conditions within the compressor section of the jet engine by oscillating and/or dithering of each engine's environmental control unit settings. This system is comprised of an anti-icing control routine 170 that governs behavior of subroutines that are nested within the program; a power ON or OFF conditional 171 that determines whether the environmental control unit oscillation anti-icing routine 170 is set to ON. When the power is OFF or returns with a NO, then the routine reverts back to the anti-icing control routine 170 unchanged. When the power is OFF or returns with a YES, then the program proceeds to an environmental control unit oscillation schedule routine 172. This routine determines the environmental control units' oscillation schedule according to the compressor stage temperatures and/or pressure gradients for each of the jet aircraft's engines. In addition, call commands 173A and 173B are directed to an engine control subroutine, Bleed to Air Control Sub One, depicted in FIG. 19. Bleed to Air Control Sub One is used to govern the environmental control unrf s oscillation for any engine that the routine is dedicated to by the anti-icing control routine 170. Similarly, call commands 174A and 174B are directed to an engine control subroutine. Bleed to Air Control Sub Two, depicted in FIG. 20. Bleed to Air Control Sub Two is used to govern the environmental control unif s oscillation for two of the plane's four available engines. Call commands 173A, 173B, 174A, and 174B all run in parallel to each other. As each subroutine operates and the total environmental control load placed upon each engine is oscillated up and down, while the net total thrust and the net total environmental control load placed upon all engines is held constant. The output of the anti-icing control routine 170 and the call commands 173A, 173B, 174A, and 174B are in communication with at least one engine control unit 179, which possesses control over at least one throttle 111, 121, 131, 141 to at least one engine 50, 60.

[188] FIG. 18 is a flow chart for a novel gas turbine jet engine environmental control unit- based anti-icing system 110, 120, 130, 140 for a four engine aircraft. The purpose of this system is to alleviate icing conditions within the compressor section of the jet engine by oscillating and/or dithering of each engine's environmental control unit settings. This system is comprised of an ant Nc ing control routine 180 that governs behavior of subroutines that are nested within the program; a power ON or OFF conditional 181 that determines whether the environmental control unit oscillation anti-icing control routine 180 is set to ON. When the power is OFF or returns with a NO, then the routine reverts back to the anti-icing control routine 180 unchanged. When the power is OFF or returns with a YES, then the program proceeds to an environmental control unit oscillation schedule routine 182. This routine determines the environmental control units' oscillation schedule according to the compressor stage temperatures and/or pressure gradients for each of the jet aircraft's engines. In addition, a call command 183 is directed to an engine control subroutine, Bleed to Air Control Sub One, depicted in FIG. 19. Bleed to Air Control Sub One is used to start the environmental control unrf s oscillation for any engine that the routine is dedicated to by the anti-icing control routine 180. Similarly, a call command 184 is directed to an engine control subroutine. Bleed to Air Control Sub Two, depicted in FIG. 20. Bleed to Air Control Sub Two is used to start the environmental control unit's oscillation for two of the plane's four available engines. The call commands 183 and 184, and a yaw control routine 185 all run in parallel to each other. The yaw control routine 185 uses the aircraft's rudder and autopilot to maintain a straight course. Note that adjusting for yaw will only be necessary if constant fuel is maintained during the oscillation of the environmental control load. Altering the temperature gradient within the engine will still occur, even at constant thrust, once the bleed valve setting has been altered. The output of the anti-icing control routine 180 and the call commands (183, 184) the yaw adjust routine 185 are in communication with at least one engine control unit 189, which possesses control over at least one throttle 111, 121, 131, 141 to at least one engine 50, 60.

[189] FIG. 19 is a flow chart for a novel subroutine, referred to as "Bleed to Air Control Sub One", that is utilized by the anti-icing systems presented in FIG. 17 and FIG. 18. This system is comprised of an initialization routine 190 that stores the engine's environmental control load settings at the outset of the routine, an environmental control unit (ECU) load routine 191 to gradually increase the environmental control load according to the compressor stage's temperature and/or pressure gradient, an ECU peak setting conditional 192 that determines whether the engine's environmental control load has increased to its peak setting, an ECU load decrease routine 193 to offset the engine's environmental load, an ECU trough setting conditional 194 that determines whether the engine's environmental load has been offset by twice the gain that resulted from the ECU load increase routine 191, an ECU load decrease routine 195 to restore the environmental load to its initial levels, an ECU restoration conditional 196 that determines whether the engine's initial environmental control load has returned to the same setting that was stored by the initialization subroutine 190, and a return command 197 to return the program counter to the calling routine.

[190] FIG. 20 is a flow chart for a novel subroutine, referred to as "Bleed to Air Control Sub Two", that is utilized by the anti-icing systems presented in FIG. 17 and FIG. 18. This system is comprised of an initialization routine 200 that stores the engine's environmental control load settings at the outset of the initialization routine 200, an ECU load decrease routine 201 to gradually decrease the environmental control load according to the compressor stage's temperature and/or pressure gradient, an ECU trough setting conditional 202 that determines whether the engine's environmental control load has decreased to its trough setting, an ECU load increase routine 203 to increase the engine's environmental load, an ECU peak setting conditional 204 that determines whether the engine's environmental load has been increased by twice the decrease that resulted from the ECU load decrease routine 201, an ECU load decrease routine 205 to decrease the environmental load to its initial levels, an ECU restoration conditional 206 that determines whether the engine's initial environmental control load has returned to the same setting stored by the initialization subroutine 200, and a return command 207 to return the program counter to the calling routine.

Illustration Group 10

[191] FIG. 21 is a flow chart for a novel gas turbine jet engine throttle-based anti-icing system 110, 120, 130, 140 for a four engine aircraft. The purpose of this system is to alleviate icing conditions within the compressor section of the jet engine by oscillating each engine's total throttle settings in unison with the other, while maintaining constant elevation by adjusting the trim settings. This system is comprised of an anti-icing control routine 210 that governs behavior of subroutines that are nested within the program; a power ON or OFF conditional 211 that determines whether the throttle oscillation anti- icing control routine 210 is set to ON. When the power is OFF or returns with a NO, then the routine reverts back to the anti-icing control routine 210 unchanged. When the power is OFF or returns with a YES, then the program proceeds to a throttle increase routine 212A that increases the engine's throttle in anticipation of the anti-icing system's subroutines; a trim adjust routine 213A that runs in parallel with 212A and adjusts the plane's trim so as to maintain constant altitude; a throttle oscillation schedule routine 214 that determines the total throttle oscillation schedule according to each compressor stage's temperature and/or pressure gradient; four parallel call commands 215A, 215B, 215C and 215D to an engine control subroutine. Throttle Oscillation Sub One, depicted in FIG.23, that starts the routine by increasing each engine's throttle in unison with the other engines; a trim adjust routine 217 that, while running in parallel to commands 215A, 215B, 215C and 215D, adjusts the plane's trim to maintain constant altitude by using the autopilot; a power ON conditional 218 that determines whether the anti-icing system is still in operation; a throttle decrease routine 212B that decreases the engine's throttle to restore the throttle to levels prior to step 212A; and a trim adjust routine 213B that runs in parallel with 212B to adjust the plane's trim so as to maintain constant altitude. At least one engine control unit 219 with control over at least one engine (50, 60) maintains communication with the anti-icing control routine 210, and the system's other routines (212, 213, 214, 215, 217) during the anti-icing system's (110, 120, 130, 140) operation. Note that adjusting the trim may not be necessary, in that adjusting the total thrust may simply oscillate the total velocity of the plane.

[192] FIG. 22 is a flow chart for a novel gas turbine jet engine throttle-based anti-icing system 110, 120, 130, 140 for a two engine aircraft. The purpose of this system is to alleviate icing conditions within the compressor section of the jet engine by oscillating each engine's throttle settings in unison with the other engine, while maintaining constant elevation by adjusting the trim settings. This system is comprised of an anti-icing control routine 220 that governs behavior of subroutines that are nested within the program; a power ON or OFF conditional 221 that determines whether the throttle oscillation anti- icing control routine 220 is set to ON. When the power is OFF or returns with a NO, then the routine reverts back to the anti-icing control routine 220 unchanged. When the power is OFF or returns with a YES, then the program proceeds to an throttle increase routine 222A that increases the engine's throttle in anticipation of the anti-icing system's other subroutines; a trim adjust routine 223A that runs in parallel with 222A and adjusts the plane's trim to maintain constant altitude; a throttle oscillation schedule routine 224 that determines the total throttle oscillation schedule according to each compressor stage's temperature and/or pressure gradient; a call command 225 to an engine control subroutine, Throttle Oscillation Sub One, depicted in FIG. 23, that governs the anti-icing throttle oscillation for the engine that it is assigned to, beginning by increasing the engine's throttle; a parallel call command 226 to engine control subroutine, Throttle Oscillation Sub One, depicted in FIG. 23, that governs the anti-icing throttle oscillation for the engine that it is assigned to, beginning by increasing the engine's throttle; a parallel trim adjust routine 227 for adjusting the plane's trim to maintain constant altitude by using the autopilot; a power ON conditional 228 that determines whether the anti-icing control routine is still in operation; a decrease throttle routine 222B that decreases the engine's throttle to restore the throttle to levels prior to step 222A; and a trim adjust routine 223B that runs in parallel with the throttle decrease routine 222B to adjust the plane's trim so as to maintain constant altitude. At least one engine control unit 229 with control over at least one engine 50, 60 maintains communication with the anti-icing control routine 220, and the system's other routines 222, 223, 224, 225, 226, 227 during the anti-icing system's 110, 120, 130, 140 operation. Note that adjusting the trim may not be necessary, in that adjusting the total thrust may simply oscillate the total velocity of the plane. [193] FIG. 23 is a flow chart for a novel subroutine, referred to as Throttle Oscillation Sub One, that is utilized by the anti-icing systems presented in FIG. 21 and FIG. 22. This subroutine is comprised of an initialization routine 230 that stores the engine's operation settings at the outset of the initialization routine 230, a throttle increase routine 231 to gradually increase the throttle according to the compressor stage's temperature and/or pressure gradient, a throttle peak setting conditional 232 that determines whether the engine's throttle has increased to its peak setting, a throttle decrease routine 233 to offset the engine's throttle, a throttle trough setting conditional 234 that determines whether the engine's throttle has been offset by twice the gain that resulted from the throttle increase routine 231, a throttle increase routine 235 to increase the throttle, a throttle restoration conditional 236 that determines whether the engine's throttle has returned to the same setting as at the beginning of the initialization subroutine 230, and a return command 237 to return the program counter to the calling routine.

[194] FIG. 24 is a flow chart for a novel subroutine, referred to as Throttle Oscillation Sub Two, that is utilized by the engine control modules presented in FIG. 21 and FIG. 22. This system is comprised of an initialization routine 240 that stores the engine's operating settings at the outset of the initialization routine 240, a throttle decrease routine 241 to gradually decrease the throttle according to the compressor stage's temperature and/or pressure gradient, a throttle trough setting conditional 242 that determines whether the engine's pressure ratio has decreased to its trough target setting, a throttle increase routine 243 to gradually restore and overshoot the lost throttle by increasing the throttle by twice the initial decreased amount, a throttle peak setting conditional 244 that determines whether the engine's throttle peak target has been reached, a throttle decrease routine 245 to decrease the throttle, a throttle restoration conditional 246 that determines whether the engine's throttle has returned to the same setting as at the beginning of the initialization subroutine 240, and a return command 247 to return the program counter to the calling routine.

Illustration Group 11

[195] FIG. 25 is a flow chart for a novel gas turbine jet engine bleed air valve-based compressor core anti-icing system 110, 120, 130, 140 for a four engine aircraft. The purpose of this system is to alleviate icing conditions within the compressor section of the jet engine by oscillating and/or dithering of each engine's bleed air valve settings to the compressor bypass. This system is comprised of an anti-icing control routine 250 that governs behavior of subroutines that are nested within the system; a power ON or OFF conditional 251A that determines whether the bleed air valve oscillation anti-icing control routine 250 is set to ON. When the power is OFF or returns with a NO, then the routine reverts back to the anti-icing control routine 250 unchanged. When the power is OFF or returns with a YES, then the program proceeds to a bleed valve oscillation schedule subroutine 252A to calculate the bleed air oscillation schedule according to the each compressor's temperature and/or pressure gradient; parallel call commands 253A and 253B to Initialization Subroutine One, depicted in FIG. 29; parallel call commands 254A and 254B to Initialization Subroutine Two, depicted in FIG. 30; a bleed valve oscillation schedule subroutine 252B to recalculate the bleed air schedule; parallel call commands 255A and 255B to the Bleed to Bypass Sub One subroutine, depicted in FIG. 27; parallel call commands 256A and 256B to the Bleed to Bypass Sub Two subroutine, depicted in FIG. 28; a conditional 251B that determines whether the bleed air anti-icing oscillation system is set to ON; parallel call commands 257A and 257B to Restoration Subroutine One, depicted in FIG. 31; and parallel call commands 258A and 258B to Restoration Subroutine Two, depicted in FIG. 32. Note that the juxtaposition of engines using offsetting bleed settings will only be necessary if constant fuel is maintained during the oscillation of the bleed valve to the bypass. At least one engine control unit 259 with control over at least one engine (50, 60) maintains communication with the anti-icing control routine 250, and the system's other routines 252, 253, 254, 255, 257, 258 during the anti-icing system's 110, 120, 130, 140 operation. Altering the temperature gradient within the engine will still occur, even at constant throttle, once the bleed valve setting has been altered.

[196] FIG. 26 is a flow chart for a novel gas turbine jet engine bleed air valve-based compressor core anti-icing system 110, 120, 130, 140 for a two engine aircraft. The purpose of this system is to alleviate icing conditions within the compressor section of the jet engine by oscillating and/or dithering of each engine's bleed air valve to the engine bypass settings, while maintaining constant direction by adjusting the plane's rudder. This system is comprised of an anti-icing control routine 260 that governs behavior of subroutines that are nested within the program; a power ON or OFF conditional 261A that determines whether the bleed air valve oscillation anti-icing control routine 260 is set to ON. When the power is OFF or returns with a NO, then the routine reverts back to the anti-icing control routine 260 unchanged. When the power is OFF or returns with a YES, then the program proceeds to a bleed valve oscillation schedule subroutine 262A to calculate the bleed air oscillation schedule according to the each compressor's temperature and/or pressure gradients; a parallel call command 263A to the bleed air valve Initialization Subroutine One, depicted in FIG. 29; a parallel call command 263B to the bleed air valve Initialization Subroutine Two, depicted in FIG. 30; a yaw adjust routine 264A operating in parallel with the initialization routines 263A and 263B by adjusting the rudder by the autopilot; a bleed air valve oscillation schedule routine 262B to recalculate the bleed air schedule; a parallel call command 265A to the Bleed to Bypass Sub One subroutine, depicted in FIG. 27; a parallel call command 265B to the Bleed to Bypass Sub Two subroutine, depicted in FIG. 28; a yaw adjust routine 264B operating in parallel to the bleed the bypass routines 265A and 265B to adjust for yaw by adjusting the rudder by using the autopilot; a power ON conditional 261B that determines whether the bleed air anti-icing oscillation control routine is set to ON; a parallel call command 266A to the Restoration Subroutine One, depicted in FIG. 31; and a parallel call command 266B to the Restoration Subroutine Two, depicted in FIG. 32; and a yaw adjust routine 264C operating in parallel with the bleed air valve restoration routines 266A and 266B to adjust for yaw by adjusting the rudder by the autopilot. At least one engine control unit 269 with control over at least one engine 50, 60 maintains communication with the anti-icing control routine 260, and the system's other routines 262, 263, 264, 265, 266 during the anti-icing system's 110, 120, 130, 140 operation.

[197] Note that adjusting for yaw will only be necessary if constant fuel is maintained during the oscillation of the bleed valve to the bypass. Altering the temperature gradient within the engine will still occur, even at constant throttle, once the bleed valve setting has been altered. Note also that the juxtaposition of engines using offsetting bleed settings will only be necessary if constant fuel is maintained during the oscillation of the bleed valve to the bypass. Altering the temperature gradient within the engine will still occur, even at constant throttle, once the bleed valve setting has been altered.

[198] FIG. 27 is a flow chart for a novel subroutine, referred to as "Bleed to Bypass Sub One", which is utilized by anti-icing systems presented in FIG. 25 and FIG. 26. This subroutine is comprised of an initialization routine 270 that stores the engine's operation settings at the outset of the initialization subroutine 270, a bleed valve close routine 271 to gradually close the bleed valve to the bypass according to the compressor stage's temperature and/or pressure gradient, a bleed valve close conditional 272 that determines whether the compressor's temperature and/or pressure gradient has increased to its peak setting, a bleed valve open routine 273 to decrease the compressor's temperature and/or pressure gradient by opening the bleed valve to the engine's bypass while maintaining constant fuel flow, a bleed valve open conditional 274 that determines whether the bleed valve to the bypass has opened to the initialization subroutine's 270 initial setting, and a return command 275 to return the program counter to the calling routine.

[199] Note that alteration of the temperature gradient within the engine will still occur, even if a constant pressure ratio and/or engine RPM is maintained by compensating with more or less fuel flow to the engine, once the bleed valve setting has been altered. Pressure ratio is defined here as the ratio of the inlet pressure of the low pressure compressor and the exit pressure of the high pressure compressor. This can be altered by maintaining constant fuel flow to the engine while simultaneously adjusting the bleed valve to the bypass. Conversely, the pressure ratio can also be maintained constant. This can be done by increasing or decreasing the fuel flow respectively in response to changing pressure conditions created by altering the bleed valve setting.

[200] FIG. 28 is a flow chart for a novel subroutine, referred to as "Bleed to Bypass Sub Two", which is utilized by the anti-icing systems presented in FIG. 25 and FIG. 26. This subroutine is comprised of an initialization routine 280 that stores the engine's operation settings at the outset of the initialization routine 280, a bleed valve open routine 281 to gradually open the bleed valve to the bypass according to the compressor stage's temperature and/or pressure gradient, a bleed valve open conditional 282 that determines whether the compressor stage's temperature and/or pressure gradient has decreased to its trough setting, a bleed valve close routine 283 to restore the compressor stage's temperature and/or pressure gradient by closing the bleed valve to the engine's bypass, a bleed valve close conditional 284 that determines whether the bleed valve to the bypass has closed to the initialization subroutine's 280 initial levels, and a return command 285 to return the program counter to the calling routine.

[201] Note that alteration of the temperature gradient within the engine will still occur, even if a constant pressure ratio and/or engine RPM is maintained by compensating with more or less fuel flow to the engine, once the bleed valve setting has been altered. Pressure ratio is defined here as the ratio of the inlet pressure of the low pressure compressor and the exit pressure of the high pressure compressor. This can be altered by maintaining constant fuel flow to the engine while simultaneously adjusting the bleed valve to the bypass. Conversely, the pressure ratio can also be maintained constant. This can be done by increasing or decreasing the fuel flow respectively in response to changing pressure conditions created by altering the bleed valve setting.

[202] FIG. 29 is a flow chart for a novel subroutine, referred to as "Bleed Air Valve Initialization Sub One", that is utilized by the anti-icing systems presented in FIG. 25 and FIG. 26. This subroutine is comprised of an initialization routine 290 that stores the engine's operation settings at the outset of the initialization routine 290, a throttle increase routine 291 to gradually increase the engine throttle according the compressor stage's temperature and/or pressure gradient, a throttle peak setting conditional 292 to determine whether the engine's throttle has increased to its peak setting, a bleed valve open routine 293 to gradually offset the gained throttle by opening the bleed valve to the bypass, and a conditional 294 to determine whether opening the bleed valve has offset twice the initial throttle gain, and a return command 295 to return the program counter to the calling routine. Refer to FIG. 27 and FIG. 28 for an explanation of the relationship between the bleed valve to the bypass setting, fuel flow, throttle, total thrust from an engine, and the alteration of the temperature within the compressor stage.

[203] FIG. 30 is a flow chart for a novel subroutine, referred to as "Bleed Air Valve Initialization Sub Two", which is utilized by the anti-icing systems presented in FIG. 25 and FIG. 26. This subroutine is comprised of an initialization routine 300 that stores the engine's operation settings at the outset of the initialization routine 300, a bleed valve open routine 301 to gradually reduce the compressor's pressure by gradually opening the bleed valve to the bypass according to the compressor stage's temperature and/or pressure gradient, a bleed air valve open conditional 302 to determine whether the bleed valve setting has decreased to its trough target setting, a bleed valve close routine 303 to gradually restore the lost compressor pressure by closing the bleed valve to the bypass, a bleed valve close conditional 304 to determine whether the closing the bleed valve has restored its original settings, a throttle increase routine 305 to gradually increase the throttle according to the compressor stage's temperature and/or pressure gradient, a throttle peak setting conditional 306 to determine if the throttle has increased to its peak setting, and a return command 307 to return the program counter to the calling routine. Refer to FIG.27 and FIG. 28 for an explanation of the relationship between the bleed valve to the bypass setting, fuel flow, throttle, total thrust from an engine, and the alteration of the temperature within the compressor stage.

[204] FIG. 31 is a flow chart for a novel subroutine, referred to as "Restoration Subroutine One", which is utilized by the anti-icing systems presented in FIG. 25 and FIG. 26. This subroutine is comprised of an initialization routine 310 that stores the engine's operating settings at the outset of the initialization routine 310, a bleed valve close routine 311 to gradually increase the compressor pressure by gradually closing the bleed valve to the bypass, a bleed valve restoration conditional 312 to determine whether the bleed valve to the bypass has been restored to its pre-oscillation setting, a throttle decrease routine 313 to gradually decrease the engine throttle, a throttle restoration conditional 314 to determine whether the throttle has been restored to the engine's pre-oscillation throttle setting, and a return command 315 to return the program counter to the calling routine. Refer to FIG. 27 and FIG. 28 for an explanation of the relationship between the bleed valve to the bypass setting, fuel flow, throttle, total thrust from an engine, and the alteration of the temperature within the compressor stage.

[205] FIG. 32 is a flow chart for a novel subroutine, referred to as "Restoration Subroutine Two", which is utilized by the anti-icing systems presented in FIG. 25 and FIG. 26. This subroutine is comprised of an initialization routine 320 that stores the engine's operation settings at the outset of the initialization routine 320, a throttle decrease routine 321 to gradually lower the throttle, a throttle restoration conditional 322 to determine whether the throttle has decreased to its pre-oscillation setting, a bleed valve open routine 323 to gradually reduce the compressor pressure by opening the bleed valve to the bypass, a bleed valve open conditional 324 to determine whether opening the bleed valve has reduced the compressor's pressure to its target trough setting, a bleed valve close routine 325 to gradually close the bleed valve to the bypass, a bleed valve restoration conditional 326 to determine if the bleed valve to the bypass settings have been fully restored to the engine's pre-oscillation settings, and a return command 327 to return the program counter to the calling routine. Refer to FIG. 27 and FIG. 28 for an explanation of the relationship between the bleed valve to the bypass setting, fuel flow, throttle, total thrust from an engine, and the alteration of the temperature within the compressor stage. niustration Group 12

[206] FIG. 33 is a flow chart for a novel gas turbine jet engine hybrid anti-icing system 110, 120, 130, 140 for a four engine aircraft. The purpose of this system is to alleviate icing conditions within the compressor section of the jet engine by oscillating and/or dithering the settings to each engine's bleed air valve to the engine bypass, as well as each engine's throttle, to create a hybrid anti-icing oscillation system. This system is comprised of an anti-icing control routine 330 that governs behavior of subroutines that are nested within the program; a power ON or OFF conditional 331 that determines whether the hybrid oscillation anti-icing control routine 330 is set to ON. When the power is OFF or returns with a NO, then the routine reverts back to the anti-icing control routine 330 unchanged. When the power is OFF or returns with a YES, then the program proceeds to a hybrid oscillation schedule subroutine 332 to calculate the hybrid oscillation schedule according to the each compressor's temperature and/or pressure gradient; parallel calls 333A and 333B to the Hybrid Oscillation Sub One subroutine, depicted in FIG. 35; and parallel calls 334A and 334B to the Hybrid Oscillation Sub Two subroutine, depicted in FIG. 36. At least one engine control unit 339 with control over at least one engine 50, 60 maintains communication with the anti-icing control routine 330, and the system's other routines 331, 332, 333, 334 during the anti-icing system's 110, 120, 130, 140 operation. Note that the juxtaposition of engines using offsetting bleed settings will only be necessary if constant fuel is maintained during the oscillation of the bleed valve to the bypass. Altering the temperature gradient within the engine will still occur, even at constant thrust, once the bleed valve setting has been altered.

[207] FIG. 34 is a flow chart for a novel gas turbine jet engine hybrid anti-icing system for a two engine aircraft. The purpose of this system is to alleviate icing conditions within the compressor section of the jet engine by oscillating and/or dithering the settings to each engine's bleed air valve to the engine bypass, as well as each engine's throttle, to create a hybrid anti-icing oscillation system, while maintaining constant direction by adjusting the plane's rudder. This system is comprised of an anti-icing control routine 340 that governs behavior of subroutines that are nested within the program; a power ON or OFF conditional 341 that determines whether the hybrid oscillation anti-icing control routine 340 is set to ON. When the power is OFF or returns with a NO, then the routine reverts back to the anti-icing control routine 340 unchanged. When the power is OFF or returns with a YES, then the program proceeds to a hybrid oscillation schedule subroutine 342 to calculate the hybrid oscillation schedule according to the each compressor's temperature and/or pressure gradients; a parallel call 343 to the to the Hybrid Oscillation Sub One subroutine, depicted in FIG. 35; a parallel call 344 to the Hybrid Oscillation Sub Two subroutine, depicted in FIG. 36; and a parallel yaw adjust routine 345 to compensate for yaw by using the autopilot to adjust the rudder. At least one engine control unit 349 with control over at least one engine 50, 60 maintains communication with the anti-icing control routine 340 and the system's other routines 341, 342, 343, 344, 345 during the anti-icing system's 110, 120, 130, 140 operation. Note that the juxtaposition of engines using offsetting bleed settings will only be necessary if constant fuel is maintained during the oscillation of the bleed valve to the bypass. Note that adjusting for yaw will only be necessary if constant fuel is maintained during the oscillation of the bleed valve to the bypass. Altering the temperature gradient within the engine will still occur, even at constant thrust, once the bleed valve setting has been altered.

[208] FIG. 35 is a flow chart for a novel subroutine, referred to as "Hybrid Oscillation One", which is utilized by the anti-icing systems presented in FIG. 33 and FIG. 34. This subroutine is comprised of an initialization routine 350 that stores the engine's operation settings at the outset of the initialization routine 350; a throttle increase routine 351 to gradually increase the throttle according to the compressor stage's temperature and/or pressure gradient; a throttle peak conditional 352 to determine whether the throttle has increased to its target peak setting; a throttle decrease routine 353 to gradually reduce the throttle; a throttle trough setting conditional 354 to determine whether the pressure ratio has been reduced to its initial setting; a bleed valve open routine 355 to gradually open the bleed valve to the bypass according to the compressor stage's temperature and/or pressure gradient; a bleed valve open conditional 356 to determine if the bleed valve open setting target has been reached for the engine; a bleed valve restoration routine 357 to gradually close the bleed valve to the bypass to the same levels stored in the subroutine's initialization routine 350; and a return command 358 to return the program counter to the calling routine.

[209] Note that alteration of the temperature gradient within the engine will still occur, even if constant throttle is maintained by compensating with more or less fuel flow to the engine, once the bleed valve setting has been altered. The pressure ratio is defined here as the ratio of the inlet pressure of the low pressure compressor and the exit pressure of the high pressure compressor. This can be altered by maintaining constant fuel flow to the engine while simultaneously adjusting the bleed valve to the bypass. Conversely, throttle can be can be maintained by compensating for pressure losses or gains. This would be done by increasing or decreasing the fuel flow respectively in response to changing pressure conditions created by altering the bleed valve setting.

[210] FIG. 36 is a flow chart for a novel subroutine, referred to as "Hybrid Oscillation Sub Two", that is utilized by the engine control modules presented in FIG. 33 and FIG. 34. This subroutine is comprised of an initialization routine 360 that stores the engine's operation settings at the outset of the initialization routine 360; a bleed valve open routine 361 to gradually open the bleed air bypass valve according to the compressor stage's temperature and/or pressure gradient; a bleed valve open conditional 362 to determine whether the bleed valve's target setting and/or compressor's pressure trough target has been reached; a bleed valve close routine 363 to gradually restore the reduced compressor pressure by closing the bleed valve to the bypass; a bleed valve restoration conditional 364 to determine whether the bleed valve settings and/or the compressor's pressure has been fully restored to the subroutine's initial settings; a throttle increase routine 365 to gradually increase the throttle according to the compressor stage's temperature and/or pressure gradient; a throttle peak setting conditional 366 to determine if the throttle has increased to its peak target setting; a throttle decrease routine 367 to gradually reduce the engine's throttle; a throttle restoration conditional 368 to determine if the throttle has been restored to the same level as at the subroutine's initialization; and a return command 369 to return the program counter to the calling routine.

[211] Again, note that alteration of the temperature gradient within the engine will still occur, even if constant throttle is maintained by compensating with more or less fuel flow to the engine, once the bleed valve setting has been altered. The pressure ratio is defined here as the ratio of the inlet pressure of the low pressure compressor and the exit pressure of the high pressure compressor. This can be altered by maintaining constant fuel flow to the engine while simultaneously adjusting the bleed valve to the bypass. Conversely, throttle can be can be maintained by compensating for pressure losses or gains. This would be done by increasing or decreasing the fuel flow respectively in response to changing pressure conditions created by altering the bleed valve setting.

Illustration Group 13

[212] FIG. 37 is a diagram representing the temperature changes to the surface temperature of an exit guide vane during a test flight encounter with ice particle meteorological conditions. [1] This exit guide vane surface temperature time series is represented alongside the turbofan's low pressure compressor rotor speed, abbreviated as Nl. The unmodified engine's low pressure rotor speed, Nl, during ice particle meteorological conditions 370 reflects the engine stability resulting from engine modifications (heat applied to the exit guide vanes), whereas the unmodified engine's low pressure rotor speed 371 reflects the deterioration of engine performance during ice particle meteorological conditions. The modified exit guide vane temperature 372 is higher than unmodified exit guides vane temperature 374. This, and the improved thermal margin 373, is a result of heat being applied to the vanes/stators. Once the exit guide vane's temperature reaches 0 ⁰C, the heat of fusion that is released when water turns to ice briefly bounces the temperature back above freezing 375, until crossing into subzero region again 376 and remaining there 377. Dlustration Group 14

[213] Fig. 38 is a flow chart depicting how a switch located in the cockpit, and therefore available to the pilot, may serve as a basis for the pilot to control the operation of the anti- icing system. It is comprised of a cockpit 380, a mechanical, hydraulic, and/or electronic switch 381, an anti-icing oscillation system 382, and an engine control unit 383 with operative control over at least one engine's 384 fuel flow and other operating parameters. The anti-icing oscillation system 382 may communicate indirectly with the engine 384 through the engine control unit 383, or it may communicate directly with the engine 384.

Additional Design Embodiments

[214] While the vast majority of commercial turbofan engines do not incorporate variable geometry into their designs, any jet engine that does utilize variable geometry can adapt this invention to include its use. Altering variable geometry in or surrounding the engine, or the engine's angle of attack, will alter the airflow gradient within the engine. Therefore, deliberately oscillating variable geometry is a viable method for oscillating a jet engine's compressor environment gradient.

[215] Also, each engine possesses an electrical generator. Because the generator's power is derived from the operation of the engine, oscillating the electrical production of the generator can also be used to oscillate a jet engine's compressor environment gradient. However, given the importance of maintaining electrical power, this method is not emphasized.

[216] In the above representations, several flow charts depict changes of the bleed air valve to the bypass being compensated for by equal and opposite changes in pressure ratios in other parallel routines. The purpose of this is to maintain constant thrust from all engines as a whole. However, these changes in thrust would only occur if fuel flow is maintained at constant levels. This is not a requirement for this system to be effective. The temperature and pressure gradient of the engine will change once the bleed valve settings have been changed, even if throttle is directed to maintain steady engine pressure ratios. Thus, adjusting for yaw will not always be necessary. Nonetheless, each approach is a possible design embodiment.

[217] Since many modifications, variations, and changes in detail can be made to the described embodiments of the invention, it is intended that all matters in the foregoing description and shown in the accompanying drawings be interpreted as illustrative and not in a limiting sense. Thus, the scope of the invention should be determined by the appended claims and their legal equivalents.

CITATION LIST NON-PATENT LITERATURE

[1] Mason, J., J.W. Strapp, and P. Chow, "Ice Particle Threat to Engines in Flight," AIAA Paper 2006-206, 2006.

[2] Lee, S., and E. Loth, Simulation of Icing on a Cascade of Stator Blades," Journal of Propulsion and Power, Vol. 24, No. 6, November-December 2008.

[3] Strapp, J.W., P. Chow, M. Maltby, A.D. Bezer, A. Korolev, I. Stromberg, and J. Hallett, "Cloud Microphysical Measurements in Thunderstorm Outflow Regions During Allied/BAE 1997 Flight Trials," 37^th AIAA Aerospace Sciences Meeting and Exhibit, Reno, NV, Jan 11-14, 1999, AIAA 99-0498.

[4] Anonymous, "Vapor Pressure," http://en.wikipedia.org/wiki/Vapor pressure. Wikipedia.com, March 24, 2010.

[5] Mason, J., "Engine Power Loss in Ice Crystal Conditions," Boeing AERO Magazine, October-December 2007.

[6] Mason, J., "The Ice Crystal Weather Threat to Engines," Boeing Corp., October 2007.

[7] Schematic for a Boeing 747, Aerospace.web.org, http://www.aerospaceweb.ore/aircraft/ietliner/b747/b747 schem Ol.jpg. image obtained from Google Images, May 2009.

[8] Image obtained from Google Images, originally created by the UK Air Accidents Investigation Branch (UAAIB). A web link and/or other reference are no longer readily available.

[9] McVey, Oliver, Pullen, Ramani, et al, "Inclement Weather & Aircraft Engine Icing," General electric Company, October 3, 2007.

[10] Veillard, X., C. Aliaga, and W. Habashi, "FENSAP-ICE Modeling of the Ice Particle Threat to Engines in Flight," SAE Technical Paper Series, January 2007, 2007-01-3323.

Claims

CLAIMSWhat is claimed is:

Claim 1. An anti-icing oscillation system (110, 120, 130, 140, 382) for use in a gas turbine jet engine (50, 60, 384) of a jet airplane, the anti-icing oscillation system (110, 120, 130, 140, 382) comprising:

means for operattvely oscillating/dithering fuel flow (117, 127, 137, 147) into at least one combustor (115, 125, 135, 145) for the prevention of icing within the gas turbine jet engine (50, 60, 384) and a resulting failure thereof.

Claim 2. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 1, wherein the means comprises:

a fuel governor (113, 123, 133, 143);

a fuel pump (114, 124, 134, 144) in fluid communications with the fuel governor (113, 123, 133, 143); and

an oscillator (112, 122, 132, 142) operatively in communications with the fuel governor (113, 123, 133, 143) for dithering the fuel flow (117, 127, 137, 147) from the fuel pump (114, 124, 134, 144) to the at least one combustor (114, 124, 133, 143).

Claim 3. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 2, further including a valve (138, 148) in communications with a throttle (111, 121, 131, 141), the fuel governor (113, 123, 133, 143), and the governor oscillator (112, 122,132, 142).

Claim 4. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 3 or 7, wherein the oscillator (112, 122, 132, 142) comprises an oscillating hydraulic piston.

Claim 5. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 3 or

13, wherein the valve (138, 148) comprises a solenoid or a ball valve device.

Claim 6. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 2 or 3 or 4 or 8, wherein the communications comprises a fluid, a hydraulic or an electrical signal.

Claim 7. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 2, further including a bleed air valve (26, 253, 254, 255, 256, 257, 258, 263, 265, 266, 271, 273, 281, 283, 293, 301, 303, 311, 323, 325, 355, 357, 361, 363), the bleed air valve (26A, 26B) being one or more valves connecting a compressor section (14A, 14B, 25A, 25B, 33, 34) to a bypass section (16), and a hydraulic oscillator (112, 122) in communications with the bypass section (16) and the one or more bleed air valves.

Claim 8. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 7 or

13, further including a switch (381) in communications with a cockpit (380) of the airplane and the oscillator (111, 121, 382).

Claim 9. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 8, wherein the switch (381) is in mechanical, hydraulic or electrical communications.

Claim 10. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 1, the oscillation system comprising:

a fuel governor (113, 123, 133, 143);

an oscillating analog circuit (112, 122) operative Iy in communications with the fuel governor (113,123) for dithering the fuel flow (117, 127) from the fuel pump (114, 124) to at least one combustor (115,125).

Claim 11. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 10, further including a switch (118, 128) in electrical communications with the throttle (111, 121, 131, 141), the fuel governor (113, 123, 133, 143), and the oscillator (112, 122, 132, 142).

Claim 12. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 11, wherein the switch (118, 128) comprises an electrical timing circuit.

Claim 13. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 2, further including a bleed air valve (26, 253, 254, 255, 256, 257, 258, 263, 265, 266, 271, 273, 281, 283, 293, 301, 303, 311, 323, 325, 355, 357, 361, 363), the bleed air valve (26A, 26B) being one or more valves connecting a compressor section (14A, 14B, 33, 34) to a bypass section (16), and an electronic oscillating circuit (112, 122) in electrical communications with the bypass section (16) and the one or more bleed air bypass valves therein.

Claim 14. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 1, the anti-icing oscillation system further comprising:

an engine control unit (159, 169, 179, 189, 219, 229, 259, 269, 339, 349, 383) for operative control over thrust from one or more jet engines (50, 60, 384); and

a compressor stage anti-icing control routine (150, 160, 170, 180, 210, 220, 250, 260, 330, 340) for disrupting an ice accretion gradient within a compressor stage (14A, 14B, 33, 34) of the one or more jet engines (50, 60, 384) by oscillating/dithering one or more environmental conditions within the one or more jet engines (50, 60, 384).

Claim 15. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 14, wherein the engine control unit (159, 169, 179, 189, 219, 229, 259, 269, 339, 349, 383) communicates to one or more throttles (110, 120, 130, 140, 382) to fluctuate a fuel flow (117, 127, 137, 147) to one or more jet engines (50, 60, 384) in an oscillating manner.

Claim 16. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 14, wherein the engine control unit (159, 169, 179, 189, 219, 229, 259, 269, 339, 349, 383) communicates to one or more bleed air valves (26, 72, 82, 253, 254, 255, 256, 257, 258, 263, 265, 266, 271, 273, 281, 283, 293, 301, 303, 311, 323, 325, 355, 357, 361, 363) to oscillate a bypass flow to one or more of the jet engines (50, 60, 384).

Claim 17. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 16, wherein a constant thrust is maintained by modulating a fuel flow (117, 127, 137, 147) to one or more combustors (115, 125, 135, 145) of one or more jet engines (50, 60, 384).

Claim 18. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 16, wherein an oscillating thrust is maintained by a constant fuel flow (117, 127, 137, 147) to one or more combustors (115, 125, 135, 145) of one or more jet engines (50, 60, 384).

Claim 19. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 14, wherein the ice accretion gradient equilibrium is disrupted by oscillating one or more environmental control bleed air valves (72, 82) of the one or more jet engines (50, 60, 384).

Claim 20. The anti-icing oscillation system (110, 120) as in Claim 14, wherein the ice accretion gradient equilibrium is disrupted by oscillating one or more electricity generation system loads of the one or more jet engines (50, 60, 384).

Claim 21. The anti-icing oscillation system (110, 120, 130, 140, 382) as in Claim 14, wherein the ice accretion gradient equilibrium is disrupted by oscillating one or more variable geometry devices of the one or more jet engines (50, 60, 384).

Claim 22. A method for the prevention of icing within a gas turbine jet engine (50, 60, 384) and a resulting failure thereof, the method using the anti-icing oscillation system (110, 120, 130, 140, 382) of Claim 1 for use in the one or more gas turbine jet engines (50, 60, 384) of a jet airplane, the method comprising the step of:

oscillating/dithering one or more operating conditions in the one or more jet engines (50, 60, 384).

Claim 23. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 22, wherein the oscillating/dithering occurs in a compressor stage (14A, 14B, 33, 34) of the one or more jet engines (50,60, 384).

Claim 24. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 22, wherein the oscillating is of a throttle (111, 121, 131, 141) of each jet engine (50, 60, 384) so as to disrupt ice accretion equilibrium therein.

Claim 25. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 22, wherein the oscillating is of a bleed air valve (26, 72, 82, 173, 174, 183, 184, 191, 193, 201, 203, 205) leading to one or more environmental control units (73, 83) of one or more of the jet engines (50, 60, 384) of the airplane so as to disrupt ice accretion equilibrium therein.

Claim 26. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 22, wherein the oscillating is of a bleed air valve (26, 72, 82, 253, 254, 255, 256, 257, 258, 263, 265, 266, 271, 273, 281, 283, 293, 301, 303, 311, 323, 325, 355, 357, 361, 363) to one or more bypasses (16, 73, 83) of one or more of the jet engines (50, 60, 384) of the airplane so as to disrupt ice accretion equilibrium therein.

Claim 27. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 22, wherein maintaining constant thrust of one or more jet engines (50, 60, 384), by maintaining constant a pressure ratio of a low and a high pressure compressors (14A, 14B), by in turn oscillating a fuel flow (117, 127, 137, 147) to one or more combustors (115, 125, 135, 145) wherein the pressure ratio changes caused by a throttle change are in an equal and opposite manner to the pressure ratio lost or gained by manipulation of a bleed air valve (26, 253, 254, 255, 256, 257, 258, 263, 265, 266, 271, 273, 281, 283, 293, 301, 303, 311, 323, 325, 355, 357, 361, 363) to a bypass (16).

Claim 28. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 27, wherein the pressure ratio is allowed to fluctuate by in turn maintaining constant fuel flow to one or more combustors (115, 125, 135, 145) as the bypass valve (26) is manipulated in a predetermined manner.

Claim 29. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 23, wherein the operating condition is an environmental condition and environmental oscillations of each jet engine (50, 60, 384) are run in parallel (153, 154, 163, 164, 173, 174, 183, 184, 215, 225, 226, 253, 254, 255, 256, 257, 258, 263, 265, 266, 333, 334, 343, 344) to the environment oscillations being directed to each jet engine (50, 60, 384).

Claim 30. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 22, wherein allowing continual operation of the oscillating/dithering of the anti-icing system (110, 120, 130, 140, 382) where upon a completion of each oscillation cycle, the oscillation system (110, 120, 130, 140, 382) is maintained in a perpetual closed loop.

Claim 31. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 30, wherein an oscillation schedule (152, 162, 172, 182, 214, 224, 252, 262, 342} is maintained constant upon the completion of each oscillation cycle.

Claim 32. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 30, wherein an oscillation schedule (152, 162, 172, 182, 214, 224, 252, 262, 342) is recalculated upon the completion of each oscillation cycle.

Claim 33. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 30, wherein an oscillation schedule (152, 162, 172, 182, 214, 224, 252, 262, 342) is recalculated pseudo-randomly so that the oscillation schedule (152, 162, 172,

182, 214, 224, 252, 262, 342) becomes dithered.

Claim 34. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 33, wherein the oscillation schedules {152, 162, 163, 172, 173, 182,

183, 214, 224, 252, 253, 255, 257, 262, 263A, 265A, 266A, 342, 343) of two jet engines (50, 51, 52, 53, 60, 61, 62, 63) complement each other, wherein the dithered oscillation schedules (152, 162, 164, 172, 174, 182, 184, 214, 224, 252, 253B, 255B, 257B, 262, 263B, 265B, 266B, 342, 344) are mirror images.

Claim 35. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 32, wherein the oscillation schedule (152, 162, 172, 182, 214, 224, 252, 262, 342) is recalculated according to an environmental gradient with the jet engine (50, 60, 384).

Claim 36. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 32, wherein changes in the oscillation schedule (152, 162, 172, 182, 214, 224, 252, 262, 342) are quantized in such a manner as to equate to discrete zones encompassing individual sections of the jet engine (50, 60, 384), the sections comprising one rotor (24, 28) and one stator (23, 27).

Claim 37. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 36, wherein an ice accretion zone being most favorable for stator (23, 27) ice accretion may be shifted to a zone (24, 28) that is less favorable for ice accretion.

Claim 38. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 23, wherein an ice accretion equilibrium with the engine compressor stage (14, 33, 34) is disrupted by oscillating an electrical system load of one or more of the jet engines (50, 60, 384).

Claim 39. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 23, wherein a throttle (111, 121, 131, 141) of all jet engines (50, 51, 52, 53, 60, 61, 62, 63) is increased temporarily at the start of the anti-icing system (110, 120, 130, 140, 382) so as to clear away any accumulated ice within the compressor (14, 33, 34).

Claim 40. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 22, further including adjusting the low pressure compressor (14A, 33) and high pressure compressor (14B, 34) pressure ratios of all jet engines (50, 51, 52, 53, 60, 61, 62, 63) in such a manner so as to achieve constant thrust or by adjusting the RPM to achieve the same result.

Claim 41. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 24, wherein the throttle (111, 121, 131, 141) of all jet engines (50, 51, 52, 53, 60, 61, 62, 63) is coordinated in such a manner as to provide asymmetric thrust.

Claim 42. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 41, wherein a yaw of the airplane is corrected by adjusting a rudder (55).

Claim 43. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 24, wherein the throttle (111, 121, 131, 141) of all jet engines (50, 51, 52, 53, 60, 61, 62, 63) is coordinated in such a manner so as to achieve oscillating total thrust from the airplane.

Claim 44. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 24, further including maintaining a nearly constant altitude by adjusting the airplane trim.

Claim 45. The method for the prevention of icing within a gas turbine jet engine (50, 60, 384) as in Claim 24, wherein anti-icing oscillation system (110, 120, 130, 140, 382) settings that could result in a stator local wet bulb temperature of about 0° C are avoided to avoid icing of the stator (23, 27).

Claim 46. A method for the preventing of accumulating ice within a core compressor stage (14, 33, 34) of a gas turbine jet engine (50, 60, 384), comprising the step of:

using forbidden throttle and/or compressor pressure settings in order to avoid engine states that are prone to icing.

Claim 47. The method for the preventing of accumulating ice as in Claim 46, wherein a wet bulb temperature of about 0° C, or its equivalent dry bulb temperature, is used as a basis to predict a location where glaciated or mixed-phase ice particle freezing will occur in the core compressor stage (14, 33, 34).

Claim 48. The method for the preventing of accumulating ice as in Claim 47, wherein a relative location of a wet bulb 0° C temperature point in the core compressor stage (14, 33, 34) is determined by taking an absolute value of a difference between temperatures at two separate locations within the core compressor stage (14, 33, 34), the absolute value of the difference between the dry bulb equivalent to the wet bulb 0° C temperature and the temperature reading taken from one of those locations, and creating a ratio of the latter divided by the former for determing the relative location.

Claim 49. The method for the preventing of accumulating ice as in Claim 48, wherein a physical distance of a wet bulb freezing point from a location of a reference temperature reading is determined by using the ratio and multiplying it by the physical distance between the two temperature reference points.

Claim 50. The method for the preventing of accumulating ice as in Claim 46, wherein a relative location of a wet bulb freezing point in the core compressor stage (14, 33, 34), or its equivalent dry bulb temperature, is determined by referring to a database of empirically derived engine temperature and pressure relationships at a given ambient air inlet (20, 30) temperature.

Claim 51. The method for the preventing of accumulating ice as in Claim 46, wherein a relative location of a wet bulb freezing temperature in the compressor stage (14, 33, 34), or its equivalent dry bulb temperature, is estimated by a mathematical model, including but not limited to a thermodynamic model.

Claim 52. The method for the preventing of accumulating ice as in Claim 46, wherein a dew point temperature is inferred to be the same as an ambient temperature during ice particle meteorological conditions.

Claim 53. The method for the preventing of accumulating ice as in Claim 46, wherein compressor operation settings that result in a stator (23, 27) local wet bulb temperature of about 0° C, or its dry bulb equivalent, are forbidden or minimized.

Claim 54. The method for the preventing of accumulating ice as in Claim 46, wherein a 0 °C wet bulb temperature location is moved off of a stator (23, 27) by altering engine operation settings, including but not limited to adjusting a fuel flow (117, 127, 137, 147), adjusting a bleed valve to a bypass (26), adjusting an environmental control load (72, 73, 82, S3), adjusting an electricity generation load, or adjusting any variable geometry inside or outside of the jet engine (50, 60, 384).

Claim 55. The method for the preventing of accumulating ice as in Claim 46, wherein a 0 °C wet bulb temperature location is moved off of a stator (23, 27) and onto an adjacent rotor (24, 28) by altering engine operation settings, including but not limited to adjusting a fuel flow (117, 127, 137, 147), adjusting a bleed valve to a bypass (26), adjusting an environmental control load (72, 73, 82, 83), adjusting an electricity generation load, or adjusting any variable geometry inside or outside of the jet engine (50, 60, 384).

Claim 56. A method of preventing simultaneous, common mode failure of all jet engines (50, 51, 52, 53, 60, 61, 62, 63) of an airplane, the method comprising the steps of:

dithering of a steady state settings of a throttle (111, 121, 131, 141) for each jet engine (50, 51, 52, 53, 60, 61, 62, 63), such that a constant thrust from each jet engine (50, 51, 52, 53, 60, 61, 62, 63) differs from that of the other jet engines (50, 51, 52, 53, 60, 61, 62, 63).