Thrust is the most obvious of the loads on engine struc-ture. It is the reason the engine is on the aircraft in the first place. Thrust is developed not only from the ex-haust gases exiting the nozzles, but also from the pres-sure distribution due to the airflow on the nacelles. It may be that in certain sections of the casings the loads will be tensile rather than compressive. The engine may be trying to pull instead of push, as it carries the net thrust out to the connection to the aircraft. In Figure 2.29 we showed examples of how thrust is carried to the aircraft for military engines. Now consider Figure 2.43 which shows the reactions between engine and pylon for the -80A engine. The forward thrust mount assembly carries axial thrust, vertical, and side loads. The swing-ing link and fixed link connections at the rear mount
STATIC STRUCTURES 2-31
24 titanium cast/hipped reinforcing brackets
Protective cover kevlar/epoxy Dry kevlar wrap
Graphite epoxy outer casing
Aluminum inner case 2219-T85T1
Aluminum honeycomb 5056 1/8 cell- .0015 wall
£ _ J L
Figure 2.41 CF6-80C2 Fwd Fan Case Construction
CF6-Rnn? Differences from -flOA
Tl reinforcement brackets on inlet flange
Fwd flange extended forward 2" Aluminum case Iniet bolt size increased from 3/8"
to 7/16*
Bond adhesive changed to give higher strength during cure cycle
Kevlar wraps reduced from 67 (-80A) to 65 (80-C)
Thread density in axial direction reduced from 17 per in. to 15 per in.
ExoeriPnce base
CF-50 steel with stiffening rfas CF6-80A graphite/kevlar/aluminum
honey comb core/steel case
Kevlar extended aft to provide root containment
I
Figure 2.42 CF6-80C2 Containment Case
2-32 STATIC STRUCTURES
Rl vertical Rt side
Figure 2.43 Interface Load Vectors
carry side and vertical load and reaction torque. Just as the thrust reaction for the Fl 10-400 is displaced from the center line of the engine and results in significant bend-ing due to axial forces, so in the CF6-80 and -50 the ax-ial forces are reacted 20 inches from the centerline. For normal thrust alone there is a bending moment of one million inch-pounds applied to the engine at the front end of the compressor. In addition to thrust forces other forces act on the engine.
Maneuver Loads are the forces and moments generated by the inertia of the mass of the engine, both rotors and stators. Figure 2.44 taken from the basic specifications for military engines, shows the basic requirement for the FI01, F110, F404, andTF34 families of engines. For a moment, imagine what the forces on the entire structure, and in particular the mounting, must be for the F110-400 Navy engine. These engines can be subjected to 9 Gs aft during catapult and as much as 10 Gs forward during ar-rested carrier landing. As the hook engages and the air-craft slams to the deck, there can be a 10 G down load tending to bend the engine between its mounts. For the internal frame structures the gyroscopic and inertia loads on the rotors are transmitted to the bearings and, hence, to the hubs of the frame and then out to the outer shells and mounts. These forces and the acceleration forces not only load the ball and roller bearings and the static struc-tures, but also tend to bend and distort the rotors them-selves.
Pressure loads are due to high differential pressures oc-curring in the rear stages of the compressor case through die combustor case and through the high pressure tur-bine structures. Lower APs occur in the forward part of the compressor, in the low pressure turbines, and even in the fan cases and ducts. While the pressures in nacelles may be modest, diameters can be very large and the hoop stresses can be quite high. In the case of inlet struc-tures or nacelles the normal AP may be only fractions of a psi. But, under burst duct or ruptured casing conditions the inlet and nacelle structures must be able to take sev-eral psi before vents open to relieve the pressure.
Thermal loading results from different temperatures in parts of a component made from the same material, dif-ferent temperatures between connected parts as in the thermal expansion of an engine relative to the pylon or strut to which it is mounted, different coefficients of ex-pansion between connected parts of different materials all at the same temperature, different coefficients of ex-pansion along with different temperatures between ele-ments made of different materials. Thermal loads and their resulting stresses are really the forces required to deflect all of the components so they will have the same dimensions between matching points. As we have looked at the various structural elements and how they have evolved, we have seen several ways in which ther-mal mismatch has been accommodated to minimize the
STATIC STRUCTURES 2-33
MIL-E-005007E
Figure 2.44 Externally Applied Forces thermal forces. Another possibility is to select materials
with low coefficients of expansion. However, the very low thermal expansion alloys often have very poor me-chanical properties and are not suitable for prime struc-tural elements. In some cases thermal mismatch can be accommodated for through the use of pivoting links with spherical bearings at the ends. Two examples of this treatment are the main mounting links between engine and aircraft and the outer links of the turbine frames on the F101 and Fl 10 families. Here the hotter core engine is permitted to move freely relative to the colder outer duct.
Unbalanced forces in new engines occur as the result of variations in the weight of airfoils, eccentricities, and the lack of squareness in the manufacture of rotor com-ponents. The normal tolerance stack-up may result in a rotor with a slightly bent center line. The resulting un-balance is reduced to acceptable limits before the engine is shipped. But, in the course of service, airfoils may be damaged by foreign objects (hopefully, we never have any domestic object damage), erosion, deterioration of the tip caps of cooled airfoils, or fatigue or stress rupture deterioration which releases a part of the airfoil. Table 2.2 summarizes the requirements for unbalance that were developed for the CF6-50 models as a result of field problems of major unbalance. A similar table
should be prepared and added to the technical require-ments for ail engines as a guide to the design engineers.
For every engine there should be some level of unbal-ance in each rotor for which the engine can operate in-definitely with no fatigue or other damage. At a high level of damage the engine should be able to operate for 30 minutes at various power levels and then be safely shut down with no fatigue or other damage. At the high-est levels of unbalance at which the engine could con-tinue to run and generate usable thrust, 30 seconds of operation should be expected until the crew can safely shut down. At the end of that operation, there may be some minor fatigue cracks in the static structure and per-haps some repairable failures of secondary components.
However, there would be no loss of components or dam-age that could cause a fire, destroy bearings, or in any way threaten the aircraft. Finally, the ultimate unbal-ance, the worst load conceivable, is the release of a full fan blade at the disk dovetail along with all of the result-ing secondary damage. We would expect major damage in the form of cracks and tears and probable destruction of bearings, but no loss of parts or failure of major load paths, no separation from the aircraft, no threat to the safety of flight, or fires. Depending on the level of dam-age, the engine would either shut itself down or would be shut down by the crew.
2-34 STATIC STRUCTURES
LARGE HIGH-BYPASS TURBOFAN (CF6-50 MODELS)
BLADE FAILURE/UNBALANCE DURATION OF OPERATION SHUTDOWN MODE ACCEPTABLE DAMAGE
A) FAN -3000 GM-IN (TIP) UNLIMITED NONE NONE
BSTR -3000 GM-IN (> 1 STG 2 A/F) HPC -200GM-iN{1/2OFTIP, STG 1) HPT -600 GM-IN (> 1/2 of STG 1) LPT -1/5 AIRFOIL, ANY STAGE
B) FAN -15,000 GM-IN (OUTER PANEL) 5 MINUTES AT T/O, CREW ACTION NONE BSTR -15,000 GM-IN (6 STG 2 AJf) 10 MINUTES AT CLIMB,
HPC -700 GM-IN (1/2 OF STG 1) 15 MINUTES AT CRUISE, HPT -1500 GM-IN (2 STAGE 2 A/F) SAFE SHUTDOWN.
LPT -8400 GM-IN (1 STG 4 A/F)
C) FAN -50,000 GM-IN (BELOW DAMPER) 15 SECONDS AT T/O, CREW ACTION MINOR DISTORTION BSTR -50,000 GM-IN (20 STG 2 A/F) 15 SECONDS AT 50% T/O AND/OR FATIGUE
HPC -3300 GM-IN (2 STG A/F) POWER; SAFE SHUTDOWN CRACKS; REPAIRS
HPT -6750 GM-IN (3 STG. 2 AIRFOILS) REQUIRED.
LPT -17,000 GM-IN (2 STG 4 A/F)
D) FAN -310,000 GM-IN NONE REQUIRED SAFE SHUTDOWN NO LOSS OF PARTS
{-6 AND -50, 2.5 BLADES) BY ENGINE OR OR FAILURE OF
BSTR. HPC, HPT - ANY GREATER THAN C) CREW MAJOR LOAD PATHS;
FOR -80A. 250,000 GM-IN (2 BLADES) LARGE CRACKS OR
FOR -80C2, 260,000 GM-IN (1.5 BLADES) TEARS ACCEPTABLE.
Table 2.2 Typical Rotor Unbalance Design Requirements Assume for the moment that the analysis has been
per-formed and we know the stresses and deflections that re-sults from the loads discussed. What are the criteria by which we measure success? How do we know whether or not our structure is adequate for the purpose?
Limit conditions for a commercial engine is defined as one which will occur once in the life of the engine and the criteria is that there shall be no permanent deforma-tion and in some cases no loss of performance. There may be a number of limit loads defined in the engine technical requirements or in the customer's product specifications. No permanent deformation could be in-terpreted as no yielding anywhere. Usually, it means that there is no measureable deformation. Yielding can occur in very small areas of high stress. The "no loss of per-formance" is a much more severe criterion, because it means that the rotors cannot move into the stators and rub open clearances and thereby degrade the perform-ance of the components. In some military specifications, limit is defined as occurring once per flight. This means that the smalt local high stress areas, which could be al-lowed to yield in a once-in-a-lifetime case, must now be considered relative to the fatigue strength of the material.
Ultimate toads are. defined as those whichmight occur once in the lifetime of all the engines in the fleet for a particular aircraft. Significant plastic deformation can occur, but no fracture or other ultimate failures are allowed that would prevent carrying the full load, such as buckling. There may be a requirement that the engine continue to generate some fraction of its thrust capability.
Fatigue capability of the structure must be evaluated for low flight cycle fatigue variation of thrust, high maneu-ver loads, pressure, and thermal loading. Repetitive ma-neuver load information is supplied by the airframer manufacturer in the form of number of exceedances per flights or operating hours. Figure 2.45 is one such typi-cal exceedance curve for vertitypi-cal G's on a military fighter. These loads when combined with engine thrust loading produce a total mount reaction fatigue load spec-trum for analysis. Figure 2.46 displays a resulting fa-tigue load spectrum for a typical military application at one mount location and direction. In some applications there are changes of power level and maneuvers which result in smaller secondary cycles. High cycle fatigue capability must be evaluated against the stresses which are induced by unbalance or frequently repetitive
STATIC STRUCTURES 2-35
100000
Normal Load Factor, Nz (g's)
Figure 2.45 Normal Load Factor Exceedances Per 1000 Hours
{ • • • • • ^ • T : [ " • ^ • • ^ • • • • ^ • i ^ H l H H
I^^^^^^^^B | ^^^^^^^^^^^^^^^^^^^^^^^^^^^H I^^^^^^^^H I 1 ^^^^^^^B^^^^^^^^^^^^^^^^^^^M I^^^^^^^^^B | i ^^|^^^^|^^^g^^^|^^|^g^^^
I^^^^^^^^H I i mmmmmmmmmmmmmmmmm
I^^^^^^^^^B | i i ™ I ™ ^ « ™ I ^ « ^ ™ ^ « B ™ B
111111111111111 rr^mr^^M
-3000-2000-1000 0 1000 2000 3000 4000 5000 6000 7000 8000 9000 10000 11000 12000 13000 Load Level (max), Ibf.
Figure 2.46 Right Axial Engine Mount Load
2-36 STATIC STRUCTURES!
flight maneuvers such as gust loading. In general, high cycle fatigue is evaluated versus the endurance limit of the material for 10' cycles or more; low cycle fatigue is evaluated against the usual S/N curves for the material.
Severe combinations of high cycle stress superimposed on low cycle stress may require the development of ma-teria! property data for the particular case.
Damage tolerance or consideration of crack growth un-der cyclic loading, determination of critical crack size, and limitation of the crack to values that can safely carry expected loads is becoming an increasingly serious con-cern for both commercial and military engines. When a crack or defect cannot be allowed to grow safely for the life of the engine, assured inspection intervals must be defined to monitor the areas of concern.
Material properties that will be used to determine suit-ability of the design must be considered in the light of the environment in which the components must perform, specifically, the operating temperature, the possibilities of oxidation, the possibilities of corrosion, and the sta-bility of the component.
How do we analyze the static structures of an engine?
What techniques are available and how realistic are they? Initial sizing can be done with simplified stick and beam computer models, as shown in Figure 2.47. The outer shell rings can be represented as simple curved
beams. The forward and aft inner rings can be treated as curved beams with adjustment to the ring properties to account for the effects of the inner and outer shells.
Straight bars or beams represent the struts. A very sim-plified model like this could be carried out on a time sharing computer program or a PC in the office. The various sketches show the forces and moments among the elements and the distortions of the inner rings for a uniform axial load applied to the hub.
The next step up in sophistication is shown in Figure 2.48 which was used in the initial analysis of the -80A turbine rear frame. The inner shells and cones are mod-eled by interconnecting plates, struts are represented as simple bars, the polygonal is modeled between the strut ends with plate elements, and the two link reactions are shown at the ends of the struts. On the left the line of ac-tion of the pivoting link and on the right the vertical and horizontal reactions of the fixed mount point are repre-sented as forces, Models such as this are used to study overall stiffness and the forces and deflections developed between the struts and the outer shell elements of the po-lygonal ring. The values used as boundary conditions for the very detailed model of the strut end fittings and mount attaching clevises are shown in Figure 2.49. The stress distributions achieved with models of this degree of refinement can generally be used for realistic initial life calculations.
Pol "o iP«
• AH ouw r _^_
nrtg
Ring radial load patterns Fwd. inner ring Aft inner ring
Frame structure model Strut and connector loading
Figure 2.47 Stick and Beam Model
STATIC STRUCTURES 2-37
* 3 dimensional redundant structural analysis model
• Deflections/stiffnesses
* General stress levels/internal load distribution Figure 2.48 Redundant Structural Analysis Model
' 3 dimensional finite element model mount/strut section CF6-80 TRF analysis model
• Detail stress levels
Figure 2.49 Finite Element Analysis Model
2-38 STATIC STRUCTURES
To study the overall structural behavior of the engine, a technique has been developed and refined in which each major component of the engine is modeled (Figure 2.50). This model of the -80A clearly shows the Keviar containment ring, the fan cases, the fan frame struts, compressor shell and combustor cases back through the compressor rear frame, combustor casing, and turbine casings of the tangentially strutted turbine frame. In the lower half of the figure, some of the shells have been cut away to show the inner structure, the frame struts, and the bearing support cones. Models like this can be used to determine the deflections of rotors and stators and the distortion of the casings and frames as they interact un-der inertia loads, pressure loads, thrust, and thermal mismatch. They are particularly helpful in conducting
"What if. . ." studies to determine ways to minimize
distortion and deflection and improve performance by reducing running clearances between rotors and stators.
Figure 2.51 shows all of the elements of a CFM 56-3 in-stallation: the inlet and fan cowls, the basic engine struc-ture, fan reverser and the core nozzle, the pyion attachment to the aircraft. These models not only permit study of the interaction of the major components of an aircraft installation; they also are suitable for use in dy-namic studies. Figure 2.52 shows the first bending mode for the CFM56-3 engine. Such use of a finite ele-ment model permits the determination of the frequencies and the mode shapes for various natural vibration modes as well as response to force vibrations. These studies en-able us to understand the dynamics of the engine and to make modifications to the design to reduce vibratory re-sponses or change frequencies and mode shapes.
Engine structural model , ^Tyy^y^Ty
Engine structural model
%.
^ L Figure 2.50 Overall Engine Structural Model
STATIC STRUCTURES 2-39
Figure 2.51 CFM56-3 3D Engine Model
Figure 2.52 CFM56-3 First Engine Bending Mode
2-tO STATIC STRUCTURES