Every aircraft has a certain operational environment, including aspects of flight and ground operations, that it is designed to serve in throughout its lifetime. For example the operational requirements of a fighter jet are much more strenuous than those of a commercial airliner. The flight regime is broadly defined by the range of different flight speeds and altitudes called the flight envelope. Within this range lies the so-called design point, which is the operational environment in which the aircraft is expected to spend most of its time in. An example plot of a typical flight envelope is shown in Figure 1. The outline of the envelope defines the limit of performance for a specific aircraft configuration. The left edge defines the minimum speed required to keep the aircraft flying at a certain altitude. The small dip in the curve at around Mach 1.0 denotes the increase in drag caused by small supersonic pockets close to the leading edge of the airfoil. Supersonic flow is inherently terminated by a shock wave that causes an increase in fluid pressure. At speeds around Mach 1.0 these shockwaves are still located on the airfoil surface and therefore exacerbate the adverse pressure gradient across the suction surface, leading to premature boundary layer separation and higher pressure drag. The top of the curve marks the region where the minimum level speed is equal to the maximum speed that can be sustained by the aircraft’s engine and structural capability. The declining curve on the right indicates the envelope where speed is limited first by the power of the engines and then by the weight of the aircraft i.e. as the aircraft speed increases so do the loads on the airframe and therefore the material required (mass) to sustain these loads. The flight envelope in Figure 1 can be drawn for any aircraft and will be different depending on the unique role e.g. commercial transport, freight, fighter, bomber etc. Today the different roles of aircraft are no longer as clear-cut since aircraft are expected to fulfil multiple roles (e.g. freight and commercial transport) for economic reasons.
Figure 1. Flight envelope of supersonic aircraft (1)
The operating environment influences the overall shape of the aircraft which can broadly be broken down into three design segments: aerodynamic shape of wings, fuselage and controlling surfaces; the choice of propulsion; and the structural layout. Naturally, for a given design point and payload there will be conflicting requirements and optimal solutions for each area individually. However, an important point to realise is that an aircraft design will only be successful if these three design factors are dealt with concurrently i.e. the optimal compromise must be found.
The commutative property is valid for the above equation i.e. the shape of the aircraft is defined by the operational requirements, similarly the shape given to an aircraft restricts the functions that the aeroplane is capable of. This means that in the early parts of the design process the engineers need to be aware what variables can be fixed and where flexibility can be maintained to limit limitations if the design environment changes.
This picture is often complicated by a additional demands that have nothing to do with the flight envelope. Thus under the given flight envelope the engineers deal with added issues of economic requirements, manufacturability, passenger ergonomics and safety, airfield requirements, environmental and noise regulation. For example, an airline operator wishes to maximise profit on each flight and therefore a major incentive for commercial aircraft manufacturers’ is to cater to this need and not the goal of engineering state-of-the-art technology. Freight and travel airlines are in the business of making money from the payload they transport from A to B. The higher the profit per kg of payload carrier the better for the airline. In this respect the dry mass of the aircraft is of critical importance for profitability. The lower the dry mass of the airline the more payload can be carried over a certain distance for a given amount of fuel. Thus not only is the fuel efficiency improved (lower costs) but the revenue is also increased by carrying more goods. This is one of the reasons why lightweight composite materials are such a big driver for future aircraft design.
References
(1) Stinton, D. The Anatomy of the Airplane. 2nd Edition. Blackwell Science Ltd. (1998).
The treatment of defects in aircraft structural design has been an important aspect in aircraft structural design during the last 50 years. Various different catastrophic events have led to key insights that now shape the design philosophy for primary aircraft structures. One of these is the distinction between Safe-Life and Fail-Safe structures. Safe-Life components are designed to go through their service life without cracks and defects playing a major role in the stress state of the component. Thus, the required fatigue life to initiate a crack is kept below the anticipated service life. This design approach is mainly used for components for which there are no back-ups in place and where failure would lead to the loss of the aircraft. A typical example of a Safe-Life component is the landing gear and this remains one of the reasons why landing gears are made from high-strength steel for which engineers have a long history of structural data. The second “Fail-Safe” design philosophy assumes that any real manufacturing process will induce defects within the part that, even if microscopic, may vary between different batches and may grow during the service life. Thus the Fail-Safe components are structurally designed to withstand all imposed loads up to a certain certain level of defect, known as the “critical size”, which can usually be detected by eye and act as stress concentrators. In this manner critical components are continuously monitored at specific service intervals to make sure that no crack exceeds the critical defect size, and is subsequently replaced if this happens. Furthermore, crack propagation analyses are employed in order to ascertain how many flights/load cycles it will take to grow a crack to the critical size. Most of these insights stem from the experience engineers have gained during the last 50 years with metal aircraft and in fact there was quite a steep learning curve during the transition years from wood to metal aircraft.
Today we are facing a similar transition from metal to mostly fibre reinforced plastics and other advanced materials whose failure mechanisms are often much more complex than that of metals. First, in metallic structures a crack typically initiates at an imperfection or stress concentration and then propagates under fatigue loading until final failure. The damage morphology in composites however is completely different: a large number of microscopic defects, such as micro-cracks that occur during post-cure shrinkage of the resin are present over a large volume of the material and these may develop into different failure mechanism over time. Second, most metals have a ductile failure mechanism such that overloading can visually be detected by the onset of plastic deformation. Therefore there is often a warning period between a structure being overloaded and failing catastrophically. Fibre reinforced plastics, especially carbon fibre composites on the hand fail by more brittle and therefore sudden mechanisms. Third, while a major driver of component design for metal structures is crack growth, which can be predicted quite accurately today using analytical methods or Finite Element codes, fibre reinforced plastics have a plethora of other failure mechanisms and manufacturing defects that are equally important. Some examples are fibre breakage, matrix cracks, matrix-fibre debonding, delaminations, voidage, misplacement of plies, lack of impregnation and fibre waviness. Interlaminar failures such as delaminations are especially important since they can occur very quickly when a laminate is loaded through the thickness, for example at stringer run-outs, in corner-radii of C-spars or simple impact events such as tool drop in the factory. Since there are typically no reinforcing fibres in the perpendicular direction the structural integrity is only guaranteed by the weak matrix. Due to this inherent weakness different plies may literally be pulled apart at their lamination interfaces. Techniques such as through thickness reinforced such as 3D braiding, Z-Pinning or nano-fibre reinforcement are currently being researched. Under compressive forces these delaminations may form blisters, so called delimitation buckling, which can easily propagate along the lamination interface leading to disintegration of the part.
Fig. 1 Delamination Buckling in Composite Laminate
Finally, different failure mechanisms actually interact making accurate predictions of the failure load including a defect extremely difficult. Furthermore, even experimental data for laboratory sized specimens cannot readily be used for real-sized components since the scaling up of structures has been found to greatly alter the dominant failure mechanism. Finally, failure sites in fibre reinforced plastics are often internal meaning that an engineer will not be able to detect them by simple visual investigation during service intervals. As a result, the increasing use of fibre reinforced plastic construction during the recent years and near future means more sophisticated evaluation techniques are required for guaranteeing safe design and operation of aircraft. Another key question is how these new types of defects can be taken into account reliably in structural design?
Compared to metallic materials composites have a very unique characteristic in that the material and structure/part are created simultaneously. This means that the amount of imperfections in the part is greatly dependent on the manufacturing process. In composite materials the fail-safe design philosophy of degrading the material properties to that including a “critical defect size” is not only important to reduce the probability of failure as in metallic structures but also because a manufacturing process free from imperfections would be financially prohibitive. Thus, the degree of process and quality control depends greatly on the safety requirements of the industry. For example, the high-volume and competitive automobile sector needs to guarantee passenger safety while keeping manufacturing costs at a minimum. In the aerospace industry however the mass of components is absolutely critical and takes precedence over the manufacturing costs. As a result the automobile industry relies more on out-of-autoclave infusion processes that allow high production volumes such as Resin Transfer Moulding, while the aerospace industry currently relies on the high-temperature, high-pressure curing environments of the autoclave that allow the manufacture of high performance parts with low, controlled level of imperfections.
Non-destructive testing (NDT) methods are often employed to detect defects inside or on the surface of a material. In general they are broken down into surface methods, bulk volume methods and global methods. These methods are typically used at the end of the manufacturing process as a quality control measure or during the life of the part to monitor and assess its fitness for continuing use. Surface methods include visual inspection techniques such as scanning the surface for obvious cracks, porosities, resin rich/starved regions or surface waviness. This is often coupled with endoscopes to examine remote or difficult to access locations. Furthermore a common technique is dye penetrant inspection where a dye is applied to external surfaces and illuminated with an ultraviolet light in order to highlight cracks on the surface that the dye has crept into. This technique was quite popular for aero engine components but is inherently quite time and labour intensive.
1. Section of material with a surface-breaking crack that is not visible to the naked eye. 2. Penetrant is applied to the surface. 3. Excess penetrant is removed. 4. Developer is applied, rendering the crack visible. (1)
Bulk volume methods range from the simple tap test to ultrasonic screening to the most sophisticated X-ray and computer tomography techniques. The choice of the method depends greatly on the type of defect that is to be detected and criticality of cycle time and production costs. Simple surface defects, core crush in sandwich structures may easily be detected using visual techniques, while tap tests can be used very effectively to determine delaminations or large internal voids. In a tap test the component is tapped lightly with a hard object such as a coin or ring which emits a very dull sound if a delimitation lies beneath the testing point. On the other hand the exact location and size of a delimitation, possible contaminations, voids or micro-porosities can only be detected with ultra-sonic or C.T. techniques. In this respect ultra-sonic scanning has developed to be the most widely-used NDT technique in the aerospace industry due to its high detection fidelity, compactness and relative low-cost compared to C.T. techniques. In ultra-sonic scanning ultrasound is projected into a component and by measuring the strength and time delay of the echo it is possible to detect inclusions (air, solid objects etc.) that differ from the host composite material.
Fig. 3. Principle of ultrasonic testing. LEFT: A probe sends a sound wave into a test material. There are two indications, one from the initial pulse of the probe, and the second due to the back wall echo. RIGHT: A defect creates a third indication and simultaneously reduces the amplitude of the back wall indication. (2)
One of the drawbacks of ultrasonic scanning is that some sort of coupling agent (typically water or a gel) is required between the probe and surface of the part to guarantee a high-quality reading. Furthermore, the scanning of large areas can be very time intensive even with the use of multi-probe ultrasonic arrays that can be rolled across a surface or controlled by a robotic arm, such that this technique is typically restricted to critical or highly-stressed components. Finally, CT techniques are currently only widely used in academia where they can give very useful insight into the exact 3D morphology of a cured part and show how and where cracks are initiated and when they propagate. Some pieces of equipment like Synchrotron radiation computed tomography at the University of Southampton can produce extremely detailed 3D plots and videos of parts under load that are very useful to help researchers understand what drives failure in composite materials.
Fig. 4. 3D Synchrotron Image (3)
Finally, in recent years global methods such as structural health monitoring have been a hot research topic. In structural health monitoring sensors such as strain gauges or fibre-bragg grating systems are embedded within the structure and provide real time data on the stress state. In this manner the health of the structure can be monitored in real time and service intervals and replacement parts be installed at the required times. However, these systems can probably not be embedded throughout an entire aircraft and require an incredible amount of storage to cope with the continual data stream.
Understanding the detrimental effects of imperfections and the damage mechanisms is essential in order to take full advantage of the benefits that high performance composites have to offer. In this respect non-destructive testing is a very valuable tool for investigating and mapping the internal condition of a component. One of the challenges facing the aerospace and automobile industries in the future is deciding what detail of non-destructive testing is required to guarantee the structural integrity of the products to a high degree of probability during the entirety of its service life and balancing this against the cost that the specific techniques incur.
In this post I want to use the sandwich panel as an example to explain some basic concepts about bending of structures. The explanations in this post are kept very basic and are similar to a first semester course in structural mechanics. Sandwich panels are an important composite structure in aerospace applications as well as in high performance automobiles, boats and wind turbines. Typically a sandwich panel is comprised of a low stiffness, low density inner core enclosed by two stiff outer skins, as shown in Figure 1, where the whole assembly is held together by some sort of structural adhesive (Figure 2). The outer skins are typically made from stiff carbon fibre or aerospace grade aluminium.
Fig. 1. A honeycomb carbon fibre sandwich panel (1)Fig. 2. Sandwich panel components and construction
The inner core is typically a Nomex or metal honeycomb, or an open or closed cell foam. Nomex is an aramid polymer similar to Nylon that is flame-resistant and can be manufactured in paper sheet form. Nomex is a great choice for the interior of aircraft cabins such as the floor panels due to its high safety in the event of fire. Multiple sheets of Nomex paper can be placed on top of each other and glued together at the node locations by lines of adhesive, which are offset spatially between different layers. This large stack of Nomex can then be cut into smaller strips and expanded to form a sheet of Nomex honeycomb. Alternatively closed cell foams such as Rohacell® are commonly used for the core, which are denser then there open cell counterparts but prevent moisture ingress in service and have better mechanical properties.
Fig. 3. Manufacture of a honeycomb sheet (2)
But what is the advantage of using a sandwich panel?
Various structures on an aeroplane are subjected to bending loads. Essentially the bending of a beam or a plate, by say some sort of pressure loading over its surface, is equivalent to grabbing the edges and applying a moment or rotation. Under pure bending Engineer’s bending theory assumes that the structure resists this moment by a linear variation of stress through its thickness. Thus, the maximum stresses occur at the top and bottom surfaces, one being compressive and the other tensile, while the stress at the middle of the beam thickness is zero. This unstressed location is called the neutral axis. For pure bending the neutral axis is always located at the centroid of the cross-section (the mid-plane for a rectangular cross-section) and can be calculated using the integral expression for the first moment of area. Therefore we can see that that the structure balances the externally applied bending moment by an internal force couple of equal magnitude where the fulcrum of the couple is the location of the neutral axis.
Fig. 4. Bending moment and internal stress distribution of beam under pure bending (3)
However this linear variation of stress is not very efficient since the cross-section of the beam is not uniformly stressed i.e. it would be more efficient if the whole cross-section was constantly loaded by the average stress to spread out the load. One method to improve the design is to cut-out the material close to the neutral axis in order to reduce structural mass as shown in Figure 5. Another possibility is to use a sandwich panel i.e. place stronger material towards the outside where it is needed and replace the interior section with a less dense and therefore lighter (and generally weaker) material to save weight.
Fig. 5. Fuselage frame with flared holes (4)
A major advantage of the sandwich construction compared to the flared hole design is that the core separates the stiff outer skins, placing them as far as possible from the neutral axis. The degree in which a structure prevents deflection in bending is known as the bending rigidity EI, where E is the Young’s modulus or stiffness of the material used and I is the second moment of area. The second moment of area I, which is the bending resistance of the cross-section, increases the more mass is located away from the neutral axis. This is analogous to rotational motion where the inertia of rotation increases the further away the rotating mass is located from the centre of rotation. In fact, as the name “second moment of area” suggests, the bending resistance increases with the square of the distance from the neutral axis. Thus a sandwich panel moves two stiff skins (high values of E such as Carbon fibre laminates) far away from the central neutral axis in order to maximise the product EI and therefore create a structure of incredibly high specific flexural stiffness i.e. high bending stiffness coupled with minimum mass. The improvements of stiffness versus weight of a sandwich panel by increasing the separation of the two face sheets is clearly illustrated in Figure 6. Here the density of the face sheets is assumed to be 15 times higher than that of the core.
Fig. 6. Stiffness vs. weight comparison for a sandwich panel
Apart from increasing the bending rigidity another advantage of using sandwich panels is that it actually concentrates the direct bending stresses (axial and shear ) in the face sheets. This is because when a structure deforms the load always distributes relative to the stiffness of the different parts. For example, when two springs are aligned in parallel and fixed on one end by a support and are displaced by the same extension x on the other end the load taken by spring 1 will be twice as high as that by spring 2 if .
Fig. 7. Two springs in parallel (5)
This is equivalent what happens to in a sandwich beam. Since the face sheets have much higher Young’s modulii than the low-density core, in bending the large majority of the direct bending loads is actually taken by the face sheets. This means that the stress distribution is no longer continuously linear through the entire cross-section as for an isotropic material in Figure 4, but actually piecewise linear and discontinuous at the interfaces. For example Figure 8 below clearly indicates how the variation of stress through the thickness of the sandwich panel changes as the stiffness mismatch between the core and face sheets is increased. As the modulus of the skins reaches 50 times that of the core there is a large jump in bending stress from just over zero to about 2 MPa. Compared to the case of equal Young’s modulus this solution is much more efficient since both the skins and the core are more uniformly stressed. The limitation of this design is that the large discontinuity of bending stress at the interface may cause excessive transverse shear stresses at the interface that can literally pull the face skins away from the core and cause de-bonding of the two parts. This is why it is important to use a core with high transverse shear modulus and strength such as honeycomb to absorb these transverse shear loads. Furthermore, the core transverse shear strength is important for resisting point or distributed pressure loadings over the surface of the face sheets and give local support for fasteners.
Fig. 8. In-plane stress profile through the thickness of a sandwich panel for various ratios of core-to-face sheet Young’s modulus
Of course there are also many drawbacks of using sandwich panels. For example when using honeycomb cores it is very hard to form complex curved shapes using the standard hexagonal matrix shape. This is because honeycomb has very high values of Poisson’s ratio such that the anti-clastic curvature effects in bending are quite pronounced. This means that when the honeycomb is bent to adhere to a certain shape it will form opposite curvature in the perpendicular direction to form a saddle shape. During in service bending deformation this will also cause the centre of the core to want to pull away from the face sheets again leading to excessive transverse shear and normal stresses at the interface and possible de-bonding of the core and face sheets. In fact de-bonding may also occur due to impact events or slow moisture ingress into the open cell honeycomb structure during service. Furthermore, when not properly designed honeycomb cores may collapse under the external pressure loading when the sandwich panel is cured in the high-temperature and pressure oven known as an Autoclave. Some of these drawbacks can be overcome by using closed-cell forms such as Rohacell®, which have lower degrees of anti-clastic curvature and, being “closed-cell”, greatly reduce the danger of water ingress into the core. The drawback of these foams is that there intrinsic higher density makes them heavier than the equivalent honeycomb solution. Alternatively, different cellular core configurations other than honeycomb such as Flex-core, rectangular and square may be used to reduce the anti-clastic curvature problem.
Fig. 9. Different cellular core styles
In metal construction the analogy to the sandwich beam is the I-beam seen in many civil constructions. Here the two flanges are located away from the neutral axis by the vertical web section. The difference in this design is that the vertical web section does also take considerable direct in-plane loads since it is of the same material and therefore stiffness as the two flanges. However, I-beams are much more cost-effective than sandwich beams since they can be easily mass-produced and do not suffer difficulties such as debonding between the face sheets and the core.
In summary a sandwich comprises,
two stiff and lightweight face sheets that predominantly take in-plane stresses and shear loads
a low-density core that takes transverse shear loads, separates the face sheets for high bending rigidity, supports the face sheets against buckling modes forming and can give local support for fastener loads
an adhesive holding the entire assembly together which transfer shear loads to the core and keeps the skins in the correct location.
There is a saying that your audience will halve for every equation you put in a piece of writing. Well, in this case I am going to make an exception and go through the detailed derivation of the Breguet Range equation. The reason for doing this is that the maths is not very difficult but the implications of the equation are known to every pilot on earth and everyone interested in flight should know about it. Simply put, the Breguet range equation tells engineers how far and airplane can fly given a certain set of parameters, and therefore greatly influences the design of modern jet engines and airframes.
A central aspect of flying further for the same amount of fuel is maximising the lift to drag ratio of your wings and airframe. Optimising this ratio gives the maximum aircraft weight (=lift at steady horizontal flight) that can be kept in the air for a given amount of engine thrust (=drag at steady horizontal flight). However, this parameter is not the primary optimum for commercial flight. Instead one wants to fly the furthest possible distance with one fuel filling. Thus to achieve the maximum possible range the quantity to be optimised is the product of flight speed (U) with lift (L) to drag (D) ratio [latex]\frac{UL}{D}[/latex]. For most long-haul journeys (~12 hours) the most time consuming part of the journey, and therefore most critical for fuel consumption is the cruise condition. During cruise conditions the band of altitudes that the airliner travels through does not vary greatly such that the local speed of sound [latex]U_s=\sqrt{\gamma R T}[/latex] where T is the local static temperature, does not vary greatly. Consequently optimising the Mach Number [latex]M=\frac{U}{\sqrt{\gamma R T}}[/latex] times the lift to drag ratio [latex]\frac{ML}{D}[/latex] is virtually the same.
Figure 1 shows experimental data of this parameter versus the lift-coefficient [latex]C_L = \frac{2L}{\rho v^2 S}[/latex] for a Boeing 747-400 at 35,000 ft. At each Mach number L/D rises to a maximum until further increase in lift coefficient leads to stall of the aerofoil. At lower flight speeds the boundary layer separation will occur naturally towards the trailing edge but as we approach a flight speed of Mach 1 shock waves also come into play. The graph shows that for all cruise speeds the optimum value of [latex]\frac{ML}{D}[/latex] occurs at a lift coefficient of about 0.5. The wing area S of an aeroplane is set largely by conditions at take-off and landing, such that it is hard to continually operate at a lift coefficient of 0.5 as the weight and therefore lift of the aeroplane decreases as fuel is burnt. To operate as close to optimum on can therefore decrease v, not very attractive, or decrease the density [latex]\rho[/latex] by flying at higher altitudes. Large airliners therefore typically start cruise at 31 000 ft or higher and then increase altitude in steps to fly at the optimum [latex]\frac{ML}{D}[/latex].
Fig. 1. Mach number x lift-drag versus lift coefficient for various flight Mach numbers (1).
The global maximum is achieved for a cruise speed of M = 0.86. Beyond this point [latex]\frac{ML}{D}[/latex] can be seen to fall precipitously. Since the air accelerates over the top surface of the aerofoil flight speeds close to Mach 1 can lead to local pockets of supersonic flow over the airflow. At some point these supersonic pockets terminate in a lambda shock wave across which the local air pressure increases to obey the law of thermodynamics. This increase in pressure exacerbates the adverse pressure gradient along the length of the aerofoil, leading to earlier boundary layer separation and an induced increase in drag. Furthermore, the separation caused by shock waves leads to buffeting and control problems. For this reason the typical Mach Number during cruise is set around 0.85.
The next time you fly you could easily check this using one of the onboard screens that display flight data. Take the formula [latex]M=\frac{U}{\sqrt{\gamma R T}}[/latex], and set U equal to flight speed in meters/second (= km/hr divided by 3.6), ratio of specific heat capacity [latex]\gamma = \frac{c_p}{c_v}=1.4[/latex], gas constant [latex]R= 287.05 J/(kg K)[/latex] and local temperature T in Kelvin = T in °C +273. Alternatively replacing all values in the equation we get [latex]M=\frac{U (km/hr)}{72.17*\sqrt{T(C) + 273}}[/latex]. Typical flight conditions are 880 km/hr at -60°C giving a Mach Number of 0.83.
The conventional measure of the amount of fuel used compared to the thrust produced is the specific fuel conusmption (SFC). The SFC is the fuel mass-flow rate divided by the thrust produced and therefore has units of kg/(Ns). At cruise, the rate of change of weight (dW/dt) is proportional to the fuel mass-flow rate , [latex]\dot{m}_f[/latex], such that,
If SFC, U and L/D are constant this expression can be integrated to give the final result,
[latex]\ln\left( \frac{W_2}{W_1} \right) = -\frac{g(SFC)}{UL/D} \Delta x \quad \text{or} \quad \Delta x = -\frac{UL/D}{g(SFC)} * \ln\left( \frac{W_2}{W_1} \right)[/latex]
where [latex]W_1 \text{ and } W_2[/latex] are the initial and final weights during cruise. This equation is known as the Breguet Range equation. We discussed before that [latex]\frac{UL}{D}[/latex] should be optimised to increase range. However, it can be noted that the range is inversely proportional to the SFC and since SFC is also a function of the flight speed U the situation is a bit more complicated. In reality the aim is to maximise the ratio of [latex]\frac{UL}{(SFC)D}[/latex]. Of course SFC also depends on the efficiency of the jet engines, which has been discussed in a series of previous posts (1,2,3). Furthermore, the structural weight is crucial forming a large part of [latex]W_2[/latex]. Finally the aerodynamic profile of the whole aircraft has to be optimised in order to reduce drag and thereby decrease the thrust F required to overcome it.
References
(1) Cumpsty, N.A. (2003). Jet Propulsion: A Simple Guide to the Aerodynamic and Thermodynamic Design and Performance of Jet Engines. Cambridge University Press.
As I described in a previous post, the efficiency of the gas turbine cycle increases as the turbine entry temperature (TET) is increased. Therefore the hotter the combustion gases that enter the first turbine stage the more specific power the jet engine can produce. Of course the TET is bounded by the metallurgical limits of the blade materials, specifically the blade root stress, the creep strain and the melting point of the blade material. The centrifugal stresses at the root increase linearly with the density of the blade material, and linearly with both the square of the rotational speed and the square of the ratio of root-to-tip radius. Creep is the continual and gradual extension of a material under constant load over time. Apart from distorting the physical dimensions and thereby reducing performance of the engine, the induced creep stresses exacerbate the centrifugal operating stresses and will therefore lead to premature failure of the material. A rule of thumb is that the blade life is halved (for a specific blade material and cooling technology) for each 10°C rise in temperature of the metal [1]. The TET has risen from about 1050K in 1944 to about 1750 in the 1994 Rolls-Royce Trent engine. This is partially due to the use of better materials such as Inconel and single-crystal metals with better creep and fatigue properties. However there is a bound to this solution since these nickel-based alloys are typically quite heavy, leading to an increase in centrifugal stresses at the root. Therefore more important in this development has been the technology of channelling of cold compressor air to cool the turbine blades. Using these advanced cooling techniques has allowed engineers to increase the TET beyond the melting point of the blade materials.
In a modern engine around 20% of the compressed air is bled off for cooling and sealing purposes for nozzle guide vanes and turbine blades [1]. This internal air system illustrated in Fig. 1 is also used to prevent the any hot mainstream gases from flowing over the heavily stressed blade-attachment discs and control tip clearances between turbine blades and casing. The stators and outer wall of the turbine flow passage use cooling air traveling from the compressor between the combustor and outer engine casing. The turbine rotor blades, disks and inner walls of the turbine flow passage use air bled from the compressor through inner passageways. Since the stators (or nozzle guide vanes) appear before the the first row of rotating blades, the first stage of stators are exposed to the highest temperatures, including local hot-spots from the combustor close by. The temperature at the first rotor stage is then somewhat decreased by dilution of the gases with cooling air, relative velocity effects and power extraction (by gas expansion causing a drop in temperature) from the turbine. In this manner the temperature reduces through each blade row.
Fig. 1. Detailed turbine cooling paths for stator and rotor stages [2]
The laws of thermodynamics require that due to combustion inefficiencies there be a pressure loss within the combustor. This means that the mainstream pressure at the first row of stators in the turbine directly after the combustor be lower than at the exit of the final stage of the compressor. It is this pressure difference that we use to drive the cooling air through the internal passageways and into the stators and blades. In this respect improvements in combustor design over the last years has been both an advantage and a disadvantage for cooling engineers. Improvements in combustor design has led to lower pressure losses within the compressor such that more force is available to drive the bled air to the hotter aft parts of the engine. On the other hand, with increasing compression ratios the air within the compressor naturally reaches higher exit temperatures (today around 900K !!! prior to combustion [1]) reducing the effect that the cooling air has on the turbine blades. Furthermore, the cooling air is expensive from an efficiency point of view since work has been done on the compressed fluid and we would ideally like to “waste” as little as possible for secondary cooling purposes. As in most case a compromise has to be struck between power output and turbine life.
Fig. 2. Evolution of turbine blade cooling technology [3]
Figure 2 illustrates the evolution of turbine blade cooling over the last decades. In the early days of the jet era convection cooling was extensively used where the rotating blade acts as a single-pass cross-flow heat exchanger. This means that the bled compressed air flows radially through cooling passages in one-direction from root to tip, driven by the pressure differences and centrifugal forces, thereby removing heat convected to the blade from mainstream gases from axially. Improvements in modern manufacturing technology means that it is now possible to create a serpentine labyrinth of cooling passages within the blade turning the system into a multi-pass heat exchanger with higher cooling capabilities. Typically these passageways also have internal ribs and fins to increase the internal whetted area available for cooling. Furthermore, the cooling air is also vented through tiny holes onto the blade aerofoil surface, especially near the leading edge. In the ideal case the cooling air emerges at low velocity, forming a protective cooling film around the blade, hence the name film cooling.
The general cooling principles outlined above can be extended and combined to different cooling techniques. Some research has been conducted on exotic techniques for turbine discs as using pre-swirl nozzles to swirl the cooling air in the direction of the rotating discs. The increase in kinetic energy reduces the effective temperature of the air when it enters the cooling ducts in the blades. However the flow and heat structures that arise in these systems give rise to complex centripetal and Coriolis accelerations leading to accelerations in excess of 10,000g ! [1] with cyclonic and anti-cyclonic currents that are very difficult to model accurately.
The turbine is at the heart of any jet engine with its primary task being to drive the compressor. As described previously without the compressor no mechanical work would be done on the fluid prior combustion and the thrust produced would only be a function of the chemical energy stored within the fuel. The hot combustion gases that enter the turbine directly after the combustion chamber are expanded across a series of vanes and stators, known as a stage, similar to the compressor. In the case of a turbine the fluid is expanded to extract useful work and therefore the pressure of the fluid falls across each turbine stage. Since the fluid is not working against an adverse (rising) pressure gradient boundary layer separation over the aerofoils of the turbine blades is not as critical such that turbine blades can have much more agressive angles of attack with respect to the flow. Consequently, the pressure ratio across a turbine stage can be much higher than across a compressor stage and it quite common for a single turbine stage to drive six or seven compressor stages. The amount of power that can be extracted from a turbine stage is tremendous and a single turbine blade (not the full rotor of blades) may contribute up to 250 bhp [1]. The biggest driver behind the progress in turbine technology since Whittle’s first engine in the 1930’s has been the development of advanced cooling methods and the use of high-temperature alloys.
Similar to compressors axial turbines seen on most modern jet airliners are more efficient than their radial counterparts at higher flow rates. However, radial turbines are still being used on modern aircraft for auxiliary power units. Figure 1 below shows a single-shaft three-stage axial turbine i.e. the three turbine stages drive all of the compressor stages through a single shaft.
Fig. 1. Triple Stage Turbine [2]
The hot gases that exist the combustion chamber and impinge on the first row of nozzle guide vanes that turn the flow into the rotating turbine blades at the optimal angle to extract the most amount of work. Each stage of vanes and blades expands the flow thereby resulting in a drop in enthalpy (total amount of energy in the combustion gases) and a transfer of work from the fluid to the turbine. For simple jet engines the overall performance of the engine is more effectively enhanced by developing the compressor stages. However for large by-pass turbofan engines turbine aerodynamic design is crucial. Figure 2 shows the velocity triangles for the flow passing through a single turbine stage. Separate turbine rows are typically placed very close together, around 20% of a blade chord [1], and the tangential velocity of the rotor blades ( is the rotational speed and the radius of the blades) is close to the local speed of sound.
Fig. 2. Velocity triangles for turbine stage [2]
The main function of the stator is not to do work but to add swirl to the flow into order to convert some of its internal heat into kinetic energy. The turbine rotor then extracts work from the flow by removing the kinetic energy associated with the swirl velocity. In the global reference frame of the engine the flow into the stator and rotor is highly unsteady and of great complexity. However, in a frame of reference fixed to a rotating blade it can be assumed to be fairly steady with sufficient accuracy. For the first row of stators (or nozzle guide vanes) the flow impinges parallel to the axial flow direction and is consequently turned through angle with respect to the axial direction by the stator. Thus the flow leaves the stator at with a velocity with respect to the stator which is equivalent to a velocity at an angle with respect to the rotating blade. At optimum design condition is equal to the angle of rotor blade. and are the relative exit speed and blade angle respectively, such that the turning angle is equal to . An important design parameter for turbine performance is the blade coefficient , which is the ratio of the total temperature drop (which is proportional to the work done) across a stage divided by the kinetic energy of the rotor.
High efficiency are achieved with lower temperature drops per stage and therefore smaller values of and lower turning angles . However large values of are required to reduce the number of stages and keep the weight of the engine down. Consequently a compromise has to be struck between optimising thermodynamic efficiency and weight.
If the high pressure of the fluid exiting the combustion chamber were expanded in a single stage a very high velocity close to 1500 m/s [1] would be produced, which due to losses associated with supersonic shock waves, would be impossible to use efficiently. Therefore the turbine stages make a series of incremental expansions resulting in flows just over the local speed of sound, which, as shown by the velocity triangles, is apparently reduced on entry to the next stage as a result of the change in reference frame. Thus the velocity triangles show that the velocity leaving the stator is high in the frame of reference appropriate to the stator but much lower when seen at the rotor entry . Similarly the velocity leaving the rotor is high in its relative frame of reference . but lower in the absolute frame appropriate to the stator . Thus each of the turbine rows takes in a flow which is almost axial down the engine and turns it towards the tangential thereby reducing the effective cross-sectional flow area, which, by conservation of momentum, must result in an increase in fluid velocity.
Turbine Stresses
The turbine inlet blades of the first stage are the most likely to determine the life of the engine since they are exerted to the highest fluid temperatures, highest rotational speeds and highest aerodynamic loads. Stresses in the rotor blades also place restrictions on the allowable blade heights and annulus flow area. The gross of the mechanical stresses arise from the centrifugal stresses of the rotating turbine and bending moments exerted by the flowing gases, which unfortunately are both maximum at the blade root. The problem of centrifugal root stress was previously discussed for compressor blades. The turbine blades are of course tuned such that none of its natural frequencies coincide with any rotational or fluid excitation frequencies so as to prevent resonant behaviour. The gas turbine produces higher specific power and thus efficiency as the turbine entry temperature (TET) of the gas exiting the combustion chamber is increased. Of course the TET is bounded by the metallurgy of the turbine blade materials. The TET has increased from around 800°C in 1940 to 1500°C in the 1994 Rolls-Royce Trent engine. This development has in part been due to better materials but more importantly through channelling of cold compressor air to cool the turbine blades.
In this high temperature environment the life of the turbine blades is limited by creep, which is the continual and gradual extension of a material under constant load over time. Apart from distorting the physical dimensions and thereby reducing performance of the engine, the induced creep stresses exacerbate the centrifugal operating stresses and will therefore lead to premature failure of the material. Under ambient temperature creep is often only a factor for elastomers and other plastics, but at higher temperatures the effects become increasingly more pronounced for metals as well. A rule of thumb is that the blade life is halved (for a specific blade material and cooling technology) for each 10°C rise in temperature of the metal [1]. In the early days of turbine technology blades were forged but later cast for better high temperature performance. It was then found that by elongating the metal crystals along the direction of the span, creating so called directionally solidified blades, resulted in further improvements in creep performance. The standard technique for high-performance blades is to cast the blade out of a single crystal as shown in Figure 3 below. Metals may deform by separate crystals slipping along grain boundaries, such that removing the grain boundaries all together results in great improvements in resisting creep deformation.
Fig. 3. The microstructure of the three different turbine blades [4].
A typical alloy used for turbine blades today is Inconel, a nickel-based alloying containing 13% chromium, 6% iron, with small amounts of manganese, silicon and copper. These metallurgical advances account for some of the improvements in driving up TET and turbine efficiency. The other very interesting and complicated technology are blade-cooling techniques. But that is a topic for another article all together.
In this post the design of jet engine compressors will be discussed leading to the definition of ballpark performance parameters. For smaller engines centrifugal (CF) compressors are used since they can handle smaller flow rates more effectively and are more compact than axial compressors. Axial compressors however have the advantage of a smaller frontal area for a given flow rate, can handle higher flow rates and generally have higher efficiencies than CF compressors. For larger turbines used on civil aircraft the most suitable compressor and turbine will be of the axial type. Early axial compressors were able to raise the pressure of the incoming area around 5-fold, while modern turbofan engines have pressure ratios in excess of 30:1.
Low pressure axial compressor scheme of the Olympus BOl.1 turbojet
Because the pressure rises in the direction of flow through the compressor there is an acute risk of the boundary layers separating on the compressor blades as they encounter this adverse pressure gradient. When this happens the performance of the compressor drops dramatically and compressor is said to stall. For this reason the compression is spread over a large number of compressor stages such that the smaller incremental increases in pressure across each stage allow engineers to obtain a large overall pressure ratio without incurring stall. A stage consists of a row of rotating blades called the rotor and a row of stationary blades called the stator. Each of these rows may consist of between 30–100 distinct blades and there may be up to 20 stages between the air inlet and compressor outlet. The role of the rotor blades is to accelerate the incoming air in order to increase the kinetic energy of the fluid. Across the stators the fluid is then decelerated and as a consequence the fluid pressure is increased. As the pressure and density increase across each stage the overall flow velocity is kept relatively constant by reducing the height of the blades from stage to stage. Thus the compressor tapers down from inlet to outlet.
In an attempt to reduce the number of compressor stages for a more compact engine, a designer’s goal is to maximise the pressure ratio across each stage. The stage pressure ratio is given by the following expression,
Where is the stage isentropic efficiency, is the total (stagnation) temperature, the rotary speed of the compressor, the axial speed of the fluid, the coefficient of latent fusion at constant pressure, and and the angle of the rotor blade leading and trailing edge relative to the axial flow direction.
Diagram of an axial flow compressor
The pressure ratio across each stage can be maximised by increasing the rotary speed of the compressor , the angle through which the fluid is turned across the rotor blades and the axial speed of the fluid through the compressor. However there is a limit on the extent of these three parameters.
1. The blade tip speed and therefore is limited by stress considerations at the root. If the fan is assumed to be of constant cross-sectional area then the centrifugal stress at the root is given by,
Where is the tip speed, is the density of the blade, and the ratio is called the root-to-tip ratio of the blade. To prevent the blades from detaching from the hub and destroying the engine this root stress is not allowed to exceed a certain proof stress. It can be seen that the root stress is proportional to the square of the compressor rotational velocity and decreases as the blade length becomes shorter. Since the first compressor blades have the highest blade lengths they limit the maximum tip speed and therefore the efficiency of the compressor. It is therefore common to split the compressor into double or triple spool configurations such as a large fan, intermediate-pressure and high-pressure compressors that are rotating at three different speeds. In this manner the large diameter fan can rotate at lower speeds to satisfy the stress restrictions while the shorter blade high-pressure compressor may rotate at higher speeds.
However the rotational speed of the fan is typically constrained by more stringent stress considerations. In a turbofan engine the large diameter fan at the front of the engine acts as a single-stage compressor. In modern turbofan engines the fan divides the flow with most of the air going to the bypass duct to a propelling nozzle and only a small portion going into the core. The high root stresses caused by the long fan blades are often exacerbated by bird strikes. For mechanical reasons a lower limit of root-to-tip ratio of 0.35 is often employed. The flow impinging onto the fan is also at a very high Mach number since the cruising speed of civil aircraft is typically around M = 0.83. Supersonic flow inevitably terminates in a shock wave with a resulting increase in pressure and entropy over the compressor blades. Shock waves reduce the efficiency of the compressor blades since they disturb the flow over the profile that lead to boundary layer separation. Furthermore, these shock waves may cause unwanted vibrations of the fan blades that further reduce the efficiency of the compressor and increase noise. Therefore for reasons of efficiency, reducing noise and limiting the damage of bird strikes the tip speed of the fan is restricted, typically a relative Mach number of 1.6 is considered as the upper limit.
2. The axial speed has to be maximised to optimise the pressure ratio and reduce the frontal area of the engine. Similar to the argument given above the axial speed is typically limited by compressibility effects of supersonic flow. As the pressure, static temperature and therefore the speed of sound increases from stage to stage, the compressibility effects are worst in the first stages. For the first stage the air enters axially such that by adding the orthogonal velocity vectors and we get where V is the speed relative to the blade. In modern engines may be in the transonic region incurring quite large losses. In this respect twin-spool engines have the advantage that the lower-pressure compressor rotates at a lower speed, which reduces the compressibility problem.
3. The angle through which the fluid is turned across the rotor blades b is limited by the growth of the boundary layers. Compressor blades are aerofoils that function in the same manner as aeroplane wings. Therefore as the angle of attack or camber of aerofoil is increased to increase the rotation of the flow velocity vector, the adverse pressure gradient across the suction surface increases, until at some point the boundary layer separates. As the boundary layer separates the effective turning angle b is reduced such that the total pressure increase across the stage reduces.
The limits of , and all place limits on the maximum pressure ratio that can be achieved in an axial compressor. Typical examples are 350 m/s, = 200 m/s, 45°.
Compressor blades are typically quite thin and are constructed from lightweight metallic alloys such as aluminium and titanium. The compressor blades feature an aerofoil section as shown in the Figure below. The centrifugal forces that act on the airflow are balanced by high-pressure air towards the tip of the blade. In order to obtain this higher pressure towards the tip the blade must be twisted from root to tip in order to change the angle of incidence on the flow, and therefore control the pressure variation over the blade.
For aircraft jet propulsion there are in general four distinct designs: the turbojet, turbofan (or bypass engine), turboprop and turboshaft. This post will address the layout and design of the two most common engines used in modern aircraft, the turbojet and turbofan, and explain how their characteristics make each engine applicable for a specific task. Specifically, two important topics are addressed. The first is the multi-shaft engine with separate low-pressure and high-pressure spools and the second is the bypass engine, in which most of the air compressed by a fan bypasses the core combustor and turbine of the engine.
In general each engine is made up of four essential components: the compressor, combustor, turbine and nozzle as shown in Figure 1. The compressor raises the pressure of the incoming air before combustion, and the turbine, which extracts work from the hot pressurised combustion products, are at the heart of the engine. The role of the power turbine is not to provide thrust but to drive the compressor. The hot pressurised combustion products are expanded through a nozzle to produce thrust. In some military turbojet engines the exhaust velocity and therefore the thrust may be increased by “afterburning” in the exhaust duct.
Figure 1. Diagram of a typical gas turbine jet engine. Air is compressed by the fan blades as it enters the engine, and it is mixed and burned with fuel in the combustion section. The hot exhaust gases provide forward thrust and turn the turbines which drive the compressor fan blades.
1.1 The Turbojet
The turbojet is the earliest form of the jet engine as developed by Sir Frank Whittle and Hans von Ohain during WWII. It is no longer used for civil aircraft but predominantly used for high-velocity propulsion in military aircraft. Figure 1 shows a cross-sectional drawing of a typical turbojet engine and illustrates the typical layout of a turbojet engine with an axial compressor driven by an axial turbine, all on the same shaft. This assembly of shaft, compressor and turbine is oftentimes referred to as a “spool”. Newer engines typically have two or three spools such that the compression and expansion process in the compressor and turbine are spread over different parts. In this manner a low-pressure (LP) compressor and LP turbine are mounted on one shaft to form the LP spool. The LP shaft passes through the inside of the hollow high-pressure (HP) shaft on which are mounted the HP compressor and HP turbine. The compressor and turbine are split into separate parts to reduce centrifugal stresses in the compressor and turbine blades, and allow different parts of the compressor and turbine to be run at different speeds in order to optimise the running efficiency.
For sustained supersonic speeds a turbojet engine remains and attractive option for aircraft propulsion. The Rolls-Royce Olympus 593 is a two-shaft example that was used to propel the Concorde to twice the speed of sound.
1.2 A Note on Efficiency:
The propulsive or Froude efficiency of a jet engine is defined by the power output divided by the rate of change of kinetic energy of the air. The kinetic energy of the air represents the power input to the system. The power output is the product of force output i.e. the thrust and the resulting air speed . Although this is an approximation this equation summarises the essential terms that define aircraft propulsion. The force required to accelerate the fluid is given by the momentum equation,
Where is the mass flow rate of the air through the engine, is the velocity of the air entering and the velocity of the air leaving the engine. Thus there will be an equal and opposite force acting on the engine called the net thrust. The term is called the gross momentum thrust and is called the ram drag. Thus, for a turbojet the power output is,
So that,
For a fixed airspeed , can be increased by reducing . However decreasing decreases the thrust unless is increased. Thus, for civil aircraft when the economy is important is increased using high by-pass ratios of the turbofan, while for military engine where thrust is important low-by pass engines with large exit velocities are employed.
1.3 Optimisation of the Turbojet
When optimising the jet engine performance two parameters are typically considered: the specific thrust (ST) of the engine, and specific fuel consumption (SFC), the mass flow rate of fuel required to produce a unit of thrust. Generally speaking turbine designer have two thermodynamic variables to optimise these two entities: the compressor pressure ratio (R) and the turbine inlet temperature (TET). The effects of these two variables on SFC and ST will be considered in turn.
ST is strongly dependent on TET and TET should be maximised in order to keep the engine as small as possible for a specific amount of thrust. However, an increase in TET will incur a larger SFC at a constant R. On the other hand a gain in ST is generally more important than the penalty of higher SFC, especially at high flight speeds where a small engine is critical to minimise weight and drag.
Increasing R always causes a reduction in SFC and hence ensuring efficient compression stages is critical for an economic engine. For a fixed value of TET increasing R will initially result in more ST but will eventually cause ST to decrease again. Thus, there exists an optimum value of R, which is the role of the engineer to ascertain. Furthermore, the optimum pressure ratio for maximum ST increases with increasing TET.
This optimisation of R and TET can of course not be separated from the mechanical design of the engine. Driving up TET requires the use of much more expensive alloys and cooled turbine blades, which invariably lead to an increase in cost, mechanical complexity or otherwise a reduction in engine life. Increasing R will require larger compressors and turbines that incur weight, cost and mechanical complexity penalties.
Finally for different flight speeds and flight altitudes the performance of the turbojet will vary since the mass flow rate and momentum drag vary with density of the air and forward speed. Gross thrust decreases considerably with increasing altitude due to the decreasing ambient density and pressure, but specific thrust may increase due to a lower engine intake temperature. SFC however is reduced for increasing altitude, a result that was calculated by Frank Whittle as an engineering student, and led to his motivation for developing the jet engine.
2.1 The Turbofan
As revealed above the high exit velocity of turbojet engines does not allow high propulsive efficiencies required for civil aircraft. To raise the propulsive efficiency a bypass engine, often known as a turbofan engine, is used.
The core of the turbofan engine is essentially the same as the turbojet featuring a compressor, combustion chamber and power turbine as shown in Figure 2. However the engine features a second turbine that drives a large fan at the front of the engine. This fan delivers air to a bypass duct that channels air to the exhaust nozzle without passing through a combustion chamber. For this reason designers often refer to the cold flow in the bypass duct and hot flow through the core. Mixing the colder bypass air with the hot exhaust gases from the core results in higher propulsive efficiencies and lower noise levels. Early bypass engines typically had bypass ratios (the mass flow rate of bypass air divided by the mass flow rate of air going through the core) of around 0.3 to 1.5. The arrangements for modern airliners are High-Bypass-Ratio (HBR) engines with a bypass ratio of 5 or even more. In the Rolls Royce RB211 and Trent families the fan is driven at low speed by one turbine, and two internal compressors driven by another two separate turbines to give a triple spool engine.
Figure 2. Schematic Diagram of Turbofan Engine
2.2 Optimisation of the Turbofan
For turbofan design engineers have four major variables to consider: the bypass ratio (BR), overall pressure ratio (OR), fan pressure ratio (FR) and TET. Similar to the turbojet high TET is required for increased thrust. As the FR is increased the thrust contributed by the cold flow is increased while that of the hot flow decreases since more power is required to drive the fan. There is an optimum value of FR for which the total thrust is a maximum. In actual fact the optimum value of FR when is a maximum automatically produces minimum SFC if OR and BR are fixed.
The propulsive efficiency rises and the SFC falls as BR is increased. For laung-haul subsonic aircraft SFC is important to reduce cost. For these engines BR is typically between 4 and 6 and OP and TET are high. Thrust is more important for military aircraft such that BR is typically reduced to 0.5 to 1. BR significantly affects the engine efficiency, appearance, size and weight of the engine. As the weight of the engine increases less payload can be added to the aircraft such that the airlines revenue falls. Second, increasing the lift produced by the wings to carry bigger engines automatically induces more drag. Finally, for practical reasons BR > 10 are not practical with current technology since it would be necessary to install a gear box between the driving power turbine and fan to allow the turbine to run faster. Such a design would most certainly require considerable development time and would probably incur a weight penalty that outweighs the benefits of increasing the BR. Thus optimisation of the engine cannot only be considered in terms of thermodynamic parameters and aircraft manufacturers ultimately decide which engine to install based on what design gives airlines the highest financial yield.
In a typical turbofan jet engine the oncoming airflow is compressed throughout a series of compressor stages, mixed with a fuel (typically kerosene) and combusted, drastically increasing pressure and temperature, and then expanded through a nozzle to provide thrust towards the rear of the aircraft. By accelerating the fluid towards aft, Newton’s Third Law implies that this impulse must be reacted by an equal and opposite force in the opposite direction, thus propelling the aircraft forward. However, modern jet engines are also capable of producing thrust in the opposing direction. How is this possible without completely changing the direction of airflow from the exhaust to the intake which would seriously damage various engine components?
Diagram of a typical gas turbine jet engine. Air is compressed by the fan blades as it enters the engine, and it is mixed and burned with fuel in the combustion section. The hot exhaust gases provide forward thrust and turn the turbines which drive the compressor fan blades.
Thrust reversal is achieved by momentarily diverting the hot exhaust gases towards the front of the aircraft or changing the propeller/compressor pitch so that the thrust produced is directed forward. Thus thrust will act against the forward direction of travel and provide a means of deceleration. Thrust reversal is used in some flight scenarios in order to,
Alleviate the stress and reduce wear on the brakes or to enable shorter landing distances. Reverse thrust can reduce the braking distance by a third or more!
Momentarily increase the braking force during emergencies or just after touchdown when the aircraft is still traveling at a high velocity and the residual aerodynamic lift is significant. Lift reduces the normal reaction force with the ground and therefore limits friction and grip on the tyres.
Rapid deceleration in flight to enable quick changes of speed. Most aircraft cannot operate thrust reversal in flight and the majority that can are propeller-driven.
Helping to push an aircraft back from a gate. A maneuver called “powerback”.
Almost everyone who has sat in a row near the wings will have heard reverse thrust in action before. Next time you land wait for the sudden high-pitched increase in engine noise just after touchdown.
The method to achieve thrust reversal varies greatly between the different types of engines:
Since the 1930s propeller-driven aircraft generate reverse thrust by changing the angle of attack of their controllable pitch propellers:
Older reciprocating engines and modern turboprop engines both have the ability to set the propeller angle to “flat pitch”. As a result the propellor airfoils produce no forward or reverse thrust, but large amounts of drag instead. This allows the engine speed to be kept at a constant speed while descending.
The classic approach is to pitch the propeller blades to a negative angle of attack in order to direct the thrust forward.
In jet engines thrust reversal is not accomplished by running the engine in reverse but by diverting the high-velocity exhaust jet blast to the front of the engines. This can be achieved in different ways:
The target-type thrust reverser: After the combustion chamber, reverser blades angle outward in order to prematurely redirect the high-speed jet radially outwards and towards the front of the engine. This construction generally gives the appearance of flower petals.
The clamshell type: Two reverser buckets are hinged at the aft of the engine, and when deployed, intrude into the exhaust of the engine. In this manner the jet blast is captured and re-oriented towards the front.
In a turbofan engine some of the air intake is not passed through the main part of the engine, but redirected along an outside channel without being combusted. This bypass duct is aptly named “cold flow” and this arrangement is used to save fuel and reduce engine noise. Furthermore, the bypass flow can also be used to channel air radially outwards and forwards to provide thrust reversal.
The three different types of thrust mechanisms explained above [1].
Youtube has some great videos showing thrust reversal in action.
Aircraft have changed enormously over the last century from the early Wright Flyer flown at Kittyhawk to the supersonic SR-71 Blackbird flown today. Of course the developments in aeronautical engineering can be broken down into separate divisions that have developed at different rates: a) the aerodynamics, b) power plant engineering, c) control, radios and navigation aids, d) airframe engineering (e.g. hydraulic/electrical systems, interior fittings etc.), and finally e) the structural design. For example, power plants have developed in two large steps separated by a series of sudden burst of ingenuity. In order to facilitate the first successful flight the Wright Brothers had to find a light yet powerful engine system. The next stride was the ingenious invention of the jet engine prior and during WWII by Sir Frank Whittle and Hans von Ohain. In between, the power output of piston engines “increased almost 200 times from 12 bhp to over 2000 bhp in just 40 years, with only a ten times increase in mass (3) “. As will be outlined in this article, the design of aerospace structures on the other hand has only made one fundamental stride forward, but this change was sufficient to change the complete design principle of modern aircraft. Today however, the strict environmental legislation and advent of the composite era may induce further leaps in structural design.
Fig. 1. A schematic drawing of the Wright Flyer (1)
Fig. 2. The modern supersonic SR-71 Blackbird (2)
1) Wire Braced Structures
If we look at the early design of aircraft such as the Wright Flyer in Figure 1 there can really be no misunderstanding of the construction style. The entire aircraft, including most notably the wings, forward and rear structures were all constructed from rectangular frames that were prevented from shearing (forming a parallelogram) or collapsing by diagonally stretched wire. There were two major innovative thoughts behind this design philosophy. Firstly, the idea that two parallel wings would facilitate a lighter yet stronger structure than a single wing, and secondly, that these two wings could be supported with two light wires rather than with a single, thicker wooden member. The structural advantage of the biplane construction is that the two wings, vertical struts and wires form a deep light beam, which is more resistant to bending and twisting than a single wing. Much like a composite sandwich beam it can be treated as two stiff outer skins for high bending rigidity connected by a lightweight “core” to provide resistance to shear and torsion.
Fig. 3. Cutaway drawing of the 1917 Sopwith Camel (3)
Fig. 4. Cutaway drawing of the 1935 Hawker Hurricane (3)
The biplane construction with wire bracing was the most notable feature of aircraft construction for much of the following years and paired nicely with lightweight materials such as bamboo and spruce (Figure 3). Wood is a composite of cellulose fibres embedded in a matrix of lignin and the early aeronautical engineers knew to take advantage of its high specific strength and stiffness. Strangely enough, after the era of metals we are now returning back to the composite roots of aircraft, albeit in a more advanced fashion. The biplane era lasted until the 1930s at which point metal was taking over as the prime aerospace material. Initially the design philosophy was not adapted to take full advantage of thin sheet metal manufacturing techniques such that wooden spars and struts were just replaced by thinner metal tubing. Consequently there remained a striking similarity in construction between a 1917 (Figure 3) and a 1931 (Figure 4) fighter. Even though some thin metal sheets were being used these components generally did not carry much load such that the main fuselage structure featured 4 horizontal longerons supported by vertical struts and wire bracing. This so called “Warren Girder” design can also be seen in some of earliest monoplane wing constructions such as the 1935 Hawker Hurricane. Aeronautical engineers were initially “unsure how to combine the new metal construction with a traditional fabric covering (3)” used on earlier aircraft. The onset of WWII meant that some safe and conservative design decisions were made to facilitate monoplane wings and the “Warren Girder” principle was directly copied to the internal framework of monoplane wings (Figure 5). These early designs were far from optimised and perfectly characterise the transition period between wire-frame structures and the semi-monocoque structures we use today.
Fig. 5. The Hawker Hurricane wing construction (3).
2) Semi-Monocoque Structures
The internal cross-bracing was initially acceptable for the early single or double seater aircraft, but would obviously not provide enough room for larger passenger aircrafts. To overcome this, inspiration was taken from the long tradition and expertise in boat building which had already been applied to construct the fuselages of early wooden flying boats. The highest standards of yacht construction at the time featured “bent wooden frames and double or triple skins…with a clear varnished finish…and presented a much more open and usable fuselage interior (3)”. The well-established boat building techniques were thus passed on to aircraft construction to produce newer aircraft with very smooth, aerodynamic profiles.
Fig. 6. Semi monocoque fuselage construction of an early wooden flying boat (4)
The major advantage of this type of construction is that the outer skin of the fuselage and wing no longer just define the shape and aerodynamic profile of the aircraft, but become an active load-carrying member of the structure as well. Thus, the structure becomes “multifunctional” and more efficient, unlike the braced fuselage which would be just as strong without the fabric covering the girders. As a consequence the whole structure is generally at a uniform and lower stress level, reducing stress concentrations and giving better fatigue life. Finally, as the majority of the material is located at the outer surface of the structure the second and polar moments of area, and therefore the bending and torsional rigidities are much increased. On the other hand, the thin-skinned construction means that compression and shear buckling become the most likely forms of failure. In order to increase the critical buckling loads the skins are stiffened by stringers and broken up into smaller sections by spars and ribs.
Fig. 7. Components of a semi monocoque wing (5)
Because the external skin is now a working part of the structure this type of construction became to be known as stressed skin or semi-monocoque, where monocoque means “shell in one piece” and “semi” is an english addition to describe the discrete discontinuities of internal stiffeners. The adoption of the semi-monocoque construction and a change from wood to metal naturally coincided since sheet metal production allowed a variety of thin skins to be easily manufactured quite cheaply, with better surface finish and superior material properties. Furthermore, metal construction was conducive to riveting which would overcome the adhesive problems of early wooden semi-monocoque aircraft such as the deHavilland Mosquito.
Fig. 8. Cutaway Drawing of the recently released A400M aircraft (6).
Figure 8 shows the typical construction of a modern aircraft. There have been numerous different structural arrangements over the past number of years but all generally feature some sort of vertical stiffener (ribs in the wings and rings in the fuselage) and longitudinal stiffener (called stringers). Over the years the main driver has been towards a) a reduction in the number of rivets by reverting to bonded assembly or ideally manufacturing separate components as a single piece and b) understanding the effects and growth of cracks under static and fatigue loading by building structures that can easily be inspected or have multiple redundancies (load paths). The design and manufacturing methods of semi-monocoque aircraft are now so automated that the development of a new aluminium, medium sized airliner “could be regarded as a routine exercise (1)”. However, the continuing legislative pressure to reduce weight and fuel consumption provides enough incentive for further development.
3) Sandwich Structures and Composite Materials
One of the major disadvantages of thin-skinned structures is their lack of rigidity under compressive loading which gives them a tendency to buckle. A sheet of paper nicely illustrates this point, since it is quite strong in tension but will provide no support under compression. One way of improving the rigidity of thin panels is by increasing the bending stiffness with the aid of external stiffeners, which at the same time break the structure up into smaller sections. The critical buckling load is a function of the square of the width of the plate over which the load is applied. Therefore skins can be made 4 times stronger in buckling by just cutting the width in half. As a wing bends upwards the main compressive loads act on the top skin along the length of the wing and therefore a large number of stringers are visible across the width.
Fig. 8. Buckling analysis of a stiffened wing panel. The stiffeners break the buckling mode shapes into smaller wavelengths that require higher energy to form compared to a single wave (7)
Another technique to provide more rigidity is sandwich construction. This generally features a very lightweight core, such as a honeycomb lattice or a foam, sandwiched between two thin yet stiff outer panels. Here the role of the sandwich core is to carry any shear loads and separate the two skins as far as possible. The second moment of area is a function of the cube of the depth and therefore the bending rigidity is greatly increased with this technique. Ideally, in this manner it would be possible to design an entire fuselage without any internal rings or stringers and the Beech Starship is an excellent example of a successful application. However, there are problems of forming honeycomb cores onto doubly curved shells since the material is susceptible to strong anticlastic curvature, forming a saddle shape when bent in one direction. Furthermore, there are problems with condensation and water ingress into the honeycomb cells and the ability to guarantee a good bond surface between the core and the outer skins. There is the possibility to use foam cores instead, but these tend to be heavier with lower mechanical properties. Perhaps the current trend is away from sandwich construction (10).
Fig. 9. A carbon fibre composite/honeycomb sandwich panel (9)
Fig. 10. The Beech Starship whose fuselage was design using sandwich construction with minimal internal bulkheads and ribs (8)
One of the major applications of honeycomb structures has been in combination with composite materials. Stiff carbon composite panels are the ideal candidate for the outer skins and the whole assembly can be co-cured together in an autoclave without having to perform any secondary bonding operations. Furthermore, the incredible specific strength and stiffness of carbon composites makes this combination an ultra lightweight yet resilient structure for aerospace applications. Indeed, we are now at the start of the “black” carbon age in commercial aircraft design. Apart from their excellent specific strength and stiffness properties composites exhibit the ability to tailor optimum mechanical properties by orientating the majority of plies in the direction of the load and allowing for less material waste during manufacture. As a result, the first generation of commercial aircraft that contain large proportions of composite parts, such as the Boeing 787 Dreamliner and Airbus A350 XWB, are planned to enter service throughout the next years.
Fig. 11. Considerable delamination leading to catastrophic failure (11)
Considerable effort has been made to mature composite technology in order to reduce manufacturing costs, guarantee reliably high quality laminates, understand the highly complex failure criteria and built hierarchical, multifunctional or self-healing structures. One of the major shortcomings is that the structural advantages of fibre-reinforced plastics must be viewed with respect to applications where the primary loads are aligned with the fibre direction. However, if a composite plate is subjected to significant out-of-plane stresses subsurface delaminations may develop between layers due to the weak through-thickness cohesive strength of the composite. These intralaminar delaminations are a significant problem as they are difficult to detect by visual inspection and may reduce the compressive strength of the laminate by up to 60%.
4) Novel Designs
With environmental legislation becoming ever so strict it is adamant that new concepts for lightweight and fuel efficient aircraft are found swiftly. Although the pressure on developing advanced composite materials is high it must be remembered that 100 years of innovation were required to reach the stage that large metal semi-monocoque structures could be manufactured in the 1940s and another 30 years to fully understand all failure criteria. Thus we may still require significant research and development before all current issues with composite materials are resolved. Apart from carbon fibre and other composites other researchers have been looking into completely redefining the shape of aircraft. Researchers at MIT have been developing the blended wing concept and NASA are exploring the technology of morphing or shape-changing aircraft, taking inspiration directly from nature.
Fig. 12. Illustration of the MIT Silent Aircraft concept (12).
Fig. 13. NASA morphing wing aircraft (13)
Whatever the final solution might look like the next 5o years in aerospace engineering will be incredibly innovative, ground-breaking and an exciting industry to be part of!
![Fig. 1. Triple Stage Turbine [2]](https://rgroh.com/wp-content/uploads/2013/04/triple-stage-turbine.jpg)
![Fig. 2. Velocity triangles for turbine stage [2]](https://rgroh.com/wp-content/uploads/2013/04/fig9VelTrianglesTurbine_web.jpg)
![Fig. 3. The microstructure of the three different turbine blades [4].](https://rgroh.com/wp-content/uploads/2013/04/img014.jpg)
