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Military Aviation Engines

It is well known in aircraft propulsion system design that it is more efficient to generate thrust by accelerating a large mass of air a small amount than by accelerating a small mass of air a large amount; propulsive efficiency increases as the ratio of exhaust velocity to free stream velocity decreases. For a turbofan engine, this can be accomplished by reducing the fan pressure ratio (FPR), which decreases the amount of fan air stream acceleration, and increasing the fan mass flow (fan size) to maintain thrust. An increase in fan mass flow for a given core engine size leads to higher bypass ratio (BPR). The desire for higher engine efficiency has resulted in the evolution of aircraft gas turbine engines from turbojets (BPR=0), to low bypass ratio, first generation turbofans (BPR=1-2), to todays high bypass ratio turbofans (BPR=5-10). It is possible that future designs will continue this trend, leading to very-high or ultra-high bypass ratio (UHB) engines.

Reduced FPR has complementary benefits in lower engine noise due to the strong relationship between noise and the velocity of the air exiting the engine. Low pressure ratio fans also typically require lower tip speeds which can result in lower fan noise. Although there are fundamental noise and efficiency benefits to lowering FPR, there are typically weight and drag penalties which can potentially offset those benefits. In addition, the larger fan diameter associated with lower FPR can lead to engine-airframe integration issues.

Gas turbine engines as continuous combustion, open Brayton cycle internal combustion engines have come to dominate as the power plants for larger, faster aircraft to essentially the exclusion of reciprocating engines, or internal, intermittent combustion engines, earlier used as power plants for these kinds of aircraft. This is largely because of the greater power-to-weight ratio of gas turbine engines versus internal combustion engines, especially in large horsepower engines, or, more appropriately, large thrust engines in which those large thrusts are provided with a relatively small, and so smaller drag, frontal area engine structures relative to reciprocating engines. Gas turbine engines generate such large thrusts for propulsion, or horsepower for engines with an output shaft, by combining large volumes of air with large amounts of fuel, and thereby form a jet of large velocity leading to the capability to provide desired speedy flights.

In addition to providing thrust, such gas turbine engines have operated integrated drive generators to generate electricity for the aircraft and for the engine electronic controls. The amount of electricity needed for these purposes in the past has tended to be relatively modest typically in the range of a few hundred kilowatts of electrical power but, with recently arriving new aircraft, exceeding a megawatt of power. However, there are some aircraft, usually for military uses, that have come to have needs for much larger amounts of electrical power either on a relative basis, the electrical power needed relative to the capability of the gas turbine engine available, or on an absolute basis with power needs significantly exceeding a megawatt. Furthermore, such demands for electrical power in military aircraft often occur at relatively high altitudes and often occur unevenly over relatively long time durations of use, that is, large peaks repeatedly occur in electrical power demand in the course of those long use durations.

Corresponding attempts to obtain the added power from the typical aircraft propulsive system, the gas turbine engine, that are needed to operate the concomitant much larger capacity electrical generators, either on a relative or absolute basis, will subtract significantly from the thrust output of the available turbine or turbines. Making up that thrust loss in these circumstances by operating such available turbine engines so as to increase the thrust output thereof causes the already relatively low fuel use efficiency during flight to decrease significantly which can severely limit the length of otherwise long duration uses, and also brings those engines closer to becoming operationally unstable.

First, in affecting engine operational stability, a gas turbine engine comes to have a smaller operational stability margin as the aircraft it powers gains altitude. This can be seen in a typical performance map of the high pressure compressor in a gas turbine engine such as the example shown in FIG. 1A in which there are plots presented in the same graph for both compressor stall lines and steady state operating lines, each given for different altitudes, these plots showing the ratio of the compressor outlet side pressure to the inlet side pressure versus a compressor fluid flow parameter. Transitions from one operating line to the next at a greater altitude are made along curves representing rotational speed increases resulting from the engine power being increased to gain that altitude.

The compressor fluid flow capacity at locations of passageways for flowing fluids in the engine is representable by this flow parameter, WT.sup.1/2/(PA), with W being the air mass flow rate in lbsm/sec (kg/s), T the temperature in .degree. R (.degree. K.), P the air pressure in lbs/in.sup.2 (Pa) and A the cross sectional area of the engine fluid passageways (for a compressor, the air passageways between the various blades and the direction setting vanes therein) in in.sup.2 (m.sup.2). This flow capacity is limited by the fixed cross sectional area geometry of those engine passageways (fixed except for a few variable compressor stages).

An atmospheric reality is that, as the aircraft with this gas turbine engine gains altitude, the air inlet pressures decrease, and so then do all of the pressures in the engine compressors, to thereby decrease P in the denominator of the above given flow parameter of any engine gas flow path passage. Also, the air temperatures decrease with altitude to thereby decrease T in the numerator of that flow parameter. However, the rate of decrease of the pressure is greater than the rate of decrease of the temperature resulting in the flow parameter increasing as the aircraft and turbine engine therein gain altitude to thereby cause the air flow into these passageways to be in effect "overstuffing" them as result of their fixed cross sectional areas.

In fact, as the aircraft altitude exceeds approximately 36,000 feet (.about.11,000 m), the atmosphere essentially ceases to get colder with increasing altitude than about the F. ( C.) reached at that altitude. This is so even though the atmospheric pressure continues to decrease with increasing altitude, and thus the apparent air flow through the compressor above that altitude as expressed in the flow parameter appears to increase even more effectively. That is, the pressure ratio value will move up along a corresponding rotational speed curve to a higher altitude operating line and the value of the flow parameter will then move to the right as the new altitude is achieved and the engine operating conditions will move along the abscissa axis of the graph in FIG. 1A leading to a further pressure ratio value increase along the new operating line. Thus, the engine is left operating on a higher altitude operating line in FIG. 1A at a higher pressure ratio on that line which in turn will leave it with less surge margin, that is, a lesser tolerance both for continued operation in such circumstances as transient excursions in the operating point of the engine and for the normal occurrence of surface erosion of compressor airfoils during engine operation.

Because the air density is decreasing with the increase in altitude (which leads to the Reynolds number characterizing the air that is flowing through the compressor to also be decreasing), the boundary layer turbulence along the compressor blades will also accordingly be decreasing. As this boundary layer turbulence decreases, the vulnerability of each set of compressor blades to have the air flow past them separate therefrom also increases. At some point, these separations will start on small regions of a number of the compressor airfoils and, as the aircraft climbs in altitude, the separated regions will increase in size to such an extent as to decrease the capability of the blades to compress the air at the very time that more of such capability is needed because of the increased flow parameter thereby, and this leads to efficiency losses growing until the compressor suddenly stalls. The stall line in FIG. 1A including such stall points thus decreases along the ordinate axis with increasing altitude (and the corresponding increase in the flow parameter through all of the engine) which is accompanied by corresponding decreases in air density and the Reynolds number characterizing that air flow.

The steady state engine operating represents the pressure increase versus flow characteristics of the engine high pressure compressor together with the back pressure faced by that compressor in its forcing air and combustion products fluids through the engine turbine. The first order determiner of the turbine's back pressure, or more precisely, its flow capacity is the first turbine stator vane flow area. The first rotor blade flow area and the turbine efficiency will be factors also, but to a lesser extent.

This turbine flow capacity is again represented by the flow parameter WT.sup.1/2/P, and so the turbine back pressure will increase as the altitude increases because of the corresponding increase in that flow parameter. This is because the increase in the flow parameter WT.sup.1/2/P, here too, leads to the fluids flowing through the turbine fluid passageways in effect "overstuffing" them with their fixed internal areas since, as described above, the atmospheric pressure P will drop more rapidly than the atmospheric temperature T as the aircraft and engine climb in altitude. This result of the change in atmospheric conditions is reflected as changes within the engine at every point. That is, conditions within the engine are a function of both the power level and the atmosphere in which it operates.

This is compounded, because, as the aircraft with the turbine engine climbs through the atmosphere, the atmospheric pressure drops at fairly constant rate but the temperature drops until reaching around F. ( C.) at about 36,000 feet (.about.11,000 m) where it then essentially stops declining altogether with increasing altitude. Thus, a turbine engine at a higher altitude is left with relatively high internal temperatures even at low pressure, and the factor T.sup.1/2 in the flow parameter is not changing at all as the aircraft with the engine further climbs and, again, there is a corresponding decrease in the Reynolds numbers for the fluid flows at the engine turbine. The result is seen in the rising engine operating line along the graph ordinate axis as the compressor must be operated at ever higher compressions to overcome this increasing turbine back pressure.

Thus, the compressor stall line moves down along the ordinate axis of that figure as the aircraft with the engine climbs, as described above, and so moves toward the rising steady state operating line thereby decreasing the stability margin, that is, the separation between the compressor operating point and the stall line. The region above the stall line, or the "stall zone" represents flow and pressure combinations that are too great to be maintained in the engine. The high pressure at the exit of the high pressure compressor cannot be maintained and at some point the flow through that compressor reverses. The basic problem is brought on by the inability of individual rows of compressor airfoils to provide the pressure rise demanded for proper operation of the other engine components.

The flow parameter above, and the well known Reynolds number in fluid mechanics are interrelated. As a result, the effects of the rate of decline of the atmospheric pressure P and of the atmospheric temperature T on that flow parameter are also reflected in corresponding changes in the Reynolds number of the fluids involved. The nature of the relationship beyond the fluid velocity common to each is first indicated through the mass flow rate definition as W=.rho.VA so that V is proportional to W/A, and, secondly, through the ideal gas law approximation as P=.rho.RT where R is the universal gas constant so that T/P is inversely proportional to the fluid density .rho..

From the foregoing, in the high pressure compressor, the operating line shift upward is primarily explained by events at the turbine, but the downward shift in this compressor map of the stall line is primarily explained by events at the compressor. The Reynolds number on each airfoil drops as the engine climbs and the frictional effects become more dominant. Higher Reynolds numbers indicates more turbulent flow that has high inherent loss but it also will more readily remain attached to the airfoils. On the other hand, as the Reynolds number drops, the flow will separate from those airfoils until the flow capacity of the stages drops to an unsustainable level thereby resulting in the surge into the stall zone.

Secondly in affecting engine operational stability, rapid shifts in the engine operation from the existing steady state operating point, such as suddenly required engine accelerations or sudden peaks occurring in electrical power demand in electrical power generators forcibly rotated by that engine, lead to transient excursions in the operating point of the engine. The engine operation point transition path of the high pressure compressor from the steady state operating line is above that of the steady state operating line during engine accelerations as indicated in that map to more closely approach a stall condition. Thus, such accelerations lead to engine operation closer to the unstable engine operation region in the graph, and similar operation transition paths are followed in connection with occurrences of electric power demand peaks. The engine fuel control reacts to such shifts by demanding a larger fuel flow rate into the turbine engine combustor to provide the necessary additional engine power to enable it to handle those shifts, shifts which often occur rather abruptly. The higher heat thereby released in the combustor represents a back pressure to the compressor which explains the resulting engine operating point transition movement, or transition path, being above the steady state operating line. Of course, engine deceleration directives or the occurrences of troughs in the electrical power demand brings the opposite result of the operating point following a transition path below the steady state operating line.

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Page last modified: 24-06-2015 20:05:30 ZULU