Tactical Aircraft Aerodynamic Integration

Transcription

1 Tactical Aircraft Aerodynamic Integration Jeffrey W. Hamstra and Brent N. McCallum Lockheed Martin Aeronautics Company, Fort Worth, TX, USA 1 Introduction 1 2 Classic Tactical Aircraft Aerodynamic Integration 1 3 System Engineering Requirements 5 4 Advanced Integration Technologies 7 5 Evolving Development Techniques 10 6 Modern Tactical Aircraft Case Studies: F-22 and F Summary 13 Symbols/Notation 13 References 13 Further Reading 14 1 INTRODUCTION Tactical aircraft propulsion aerodynamic integration is defined as the integration of jet engine air inlet and exhaust systems, as opposed to the physical/functional integration of the engine itself. For tactical aircraft (also known as combat or fighter/attack aircraft), aerodynamic integration demands very close coordination between the air system contractor (ASC) and propulsion system contractor (PSC). Propulsion aerodynamic integration state-of-the-art has evolved significantly since design of legacy transonic aircraft currently in the military force structure, exemplified in the United States by the F-14, F-15, F-16, and F-18. New air system-level requirements derived from a systems engineering approach, in particular, survivability and affordability, must now be considered on at least an equal basis as traditional aero-performance parameters such as installation losses, net thrust, operability, maneuverability, and maximum design speed. To satisfy evolving requirements in a multi-disciplinary design environment, new inlet and exhaust technologies have emerged, including stealth-compliant highly integrated shaping and multi-axis thrust vectoring. Further technology advancements, including structural integration and active flow control, are currently in development. These technologies have been enabled in part by the growth in computational power available to both design and test communities. Emerging combat air systems such as the F-22 and F-35 characterize the advanced inlet/exhaust designs driven by these new requirements, technologies, and development techniques. This section addresses aerodynamic integration of tactical aircraft jet engine inlet and nozzle systems from an air vehicle system-level viewpoint. For more fundamental information on jet propulsion inlet and exhaust systems, refer to Gas Turbine Engines: Inlets, Gas Turbine Engines: Nozzles, Choosing and Sizing the Propulsion System, and Basic Principles: Gas Turbine Compatibility. 2 CLASSIC TACTICAL AIRCRAFT AERODYNAMIC INTEGRATION Tactical aircraft are generally defined as those military aircraft employing weapons directly in support of ground troops, and include fighters and attack aircraft. Their performance requirements are characterized by high subsonic cruise speed, supersonic dash, air-to-air combat (which requires high

2 2 Propulsion Integration Figure 1. F-15 Propulsion aerodynamic integration features. Left part of this figure is reproduced by permission of Mike Freer Touchdown- Aviation. Right part of this figure is reproduced by permission of Lee Jangsev-Korea Air Photos. maneuverability), survivability (i.e., stealth), moderate range (1000 to 2000 km), and a wide range of payload. With the exception of hypersonic vehicles, tactical aircraft propulsion system aerodynamic integration requirements are more demanding than for any other class of aircraft. For legacy systems such as the F-14, F-15, F-16, and F-18, inlet and exhaust integration were driven primarily by aeroperformance requirements, with weight an important but secondary consideration. These aircraft feature either two-dimensional (F-14 or F-15) or conformal (F-16 and F-18) inlets, and axisymmetric, non-vectoring enginemounted exhaust systems. Engine inlet and nozzle designs of this era were based on geometry that could be analyzed with simple textbook equations and empirical design guidelines. Example design guidelines are provided in Crosthwait (1967) for inlets and Gamble, Terrell and DeFrancesco (2004) for nozzles, and a thorough historical perspective for inlet design is provided by Sobester (2007). Huenecke (1987) gives an overview of both inlet and nozzle integration for the F-15 and F-16. Figures 1 and 2 illustrate key propulsion integration features for these two aircraft. Propulsion integration has more recently evolved to a balancing of aerodynamic and survivability needs based on air system-level performance requirements. In the case of the US F-117 Stealth Fighter, survivability was the primary design driver and propulsion aerodynamics was secondary. Survivability characteristics, which include alignment of inlet and nozzle edges with vehicle planform edges, 2-D nozzles, duct shaping, and full obscuration of the engine, complicate the propulsion system aerodynamic integration. Recently, systems engineering derived requirements, including weight and life cycle cost (LCC), have played a primary role in a multi-disciplinary design environment for inlet and exhaust integration. Figure 2. F-16 Propulsion aerodynamic integration features.

3 Tactical Aircraft Aerodynamic Integration Inlet integration Tactical aircraft typically spend most of their life at subsonic conditions. For instance, based on historical usage data, the F-16 exceeds Mach 1 less than 1% percent of flight time. Subsonic inlet aerodynamic design is driven primarily by duct losses, form and spillage drag, and cowl lip separation at static or takeoff conditions. The requirement for supersonic dash, even though it may rarely occur, can increase design complexity substantially, and therefore must be considered in the propulsion aerodynamic integration. Above Mach 1.4, shock losses and spillage drag begin to dominate the inlet performance. Fixed-geometry inlets, as with the F-16 and F-18, can have adequate pressure recovery at the aerodynamic interface plane (AIP, the analytical interface between the inlet and engine) and low enough spillage drag to achieve limited supersonic flight conditions. Higher Mach or extended supersonic requirements may lead to variable geometry multi-ramp inlets, such as the 2-D F-14 and F-15 inlets or the conical F-111 translating 1/4 cone inlet. These variable geometry and multi-ramp designs are mechanically scheduled to reduce shock losses (thereby improving pressure recovery) and spillage drag across a large range of supersonic flight conditions. Variable geometry can also be used to improve inlet performance at maneuver conditions. Cawthon, Truax and Steenken (1973) provide a general approach for optimizing an inlet design with these considerations, and Imfeld (1974) provides a specific example for the F-15 inlet Inlet sizing Inlet size is another key design parameter that affects cowl lip losses, duct losses, and form and spillage drag. Generally, the inlet minimum flow area (known as the throat ) is usually sized for a Mach number of 0.65 to 0.75 at the maximum inlet airflow condition. While larger throat area is desired to minimize pressure loss and distortion, and to provide airflow growth margin, smaller throat area leads to designs that are physically smaller and easier to integrate with lower drag and lower weight. While the original F-16 inlet was designed for a throat Mach number of 0.72, subsequent aircraft versions incorporated a higher airflow engine, and a growth inlet was sized for near critical operation (M = 1) at maximum engine airflow (see Hagseth, 1987). Trade studies are required to optimize inlet size and balance pressure losses, spillage drag, consideration for distortion and engine/inlet compatibility, and physical integration. Additional discussion of inlet sizing is provided in Gas Turbine Engines: Inlets Boundary layer control Legacy tactical aircraft inlets incorporate a boundary layer diverter as identified on the F-15 in Figure 1. The boundary layer diverter offsets the inlet from the fuselage and provides a passage for the forebody boundary layer to spill between the inlet and forebody. The diverter prevents the low energy boundary layer from entering the inlet and degrading aeroperformance, and, during supersonic flight, it isolates the inlet shock from the boundary layer to improve aerodynamic stability. Tactical aircraft inlets may also contain variable bleed and bypass systems. The bleed system prevents significant boundary layer buildup on compression ramps to improve aerodynamic stability and aeroperformance. Bypass systems may be necessary to match inlet airflow to engine demanded airflow. At high supersonic speeds, engine demanded airflow decreases to prevent exceeding engine internal aerodynamic and/or material temperature limitations. However, inlet characteristic airflow may be significantly greater than the demanded airflow. A bypass system provides a means to spill the excess airflow efficiently without impacting the inlet aerodynamic stability. Bleed and bypass systems typically exhaust the air on the upper surface of the vehicle in a location with a favorable local pressure. On occasion, diverter, bleed, or bypass airflow is utilized as a secondary air source for heat exchangers, ejector nozzles, or other such functions Cowl integration Tactical aircraft inlet cowl edges are often characterized by a relatively sharp profile to minimize supersonic drag. In some cases the edges may also be swept. While sharp edges are desired for supersonic flight, they create additional performance penalties at static or low-speed conditions. Typically, a vortex will form and be ingested along the cowl lip at high engine airflow and static conditions. The vortex strength is a function of inlet sizing, lip cross-section shape, and lip sweep. These design parameters can be adjusted to mitigate the vortex, but this may also affect other considerations such as spillage drag. Other approaches to mitigate lip vortex effects include tailoring the engine control system to tolerate static distortion characteristics or the use of actuated or blow-in auxiliary inlets, which can be very useful in reducing inlet distortion and increasing pressure recovery Inlet duct integration The internal subsonic duct length of an inlet is derived from inlet and engine relative placement, and the overall internal arrangement of the aircraft. Duct lengths are characterized

4 4 Propulsion Integration by L/D (length to engine face diameter ratio) and typically range from 4 to 7. Long, straight ducts of gradual cross sectional shape and area change provide the best balance of aero-performance parameters. However, to navigate around other systems such as the landing gear, crew station, internal weapons bays, and fuel tanks, the duct may be serpentine in nature, with substantial offset in the vertical and/or horizontal planes. Rapid turning can introduce adverse pressure gradients, pressure recovery loss, or spatial distortion increase due to boundary layer growth or flow separation Inlet maneuverability Maneuverability is driven by air-to-air combat resulting in extreme attitudes, which introduces severe pressure recovery and spatial distortion challenges for the inlet system. A typical tactical aircraft may maneuver up to 30 (and more in some cases) angle-of-attack and 15 angle-of-sideslip at subsonic conditions. As speed increases, the maneuver requirements will decrease but could be up to 20 angleof-attack and 5 angle-of-sideslip even at supersonic speeds. Since these conditions are transient, distortion effects are usually the first consideration for engine/inlet compatibility, and recovery effects on thrust are secondary. Extensive guidelines for inlet and forebody integration for maneuverability considerations were developed during the 1970s and can be found in Cawthon, Truax and Steenken (1973). In some cases, such as the F-16 or Eurofighter Typhoon (see Philpot, 2000), the inlet is located below the forebody, which provides favorable local flow conditions at high angleof-attack. However, this under-fuselage location may not provide beneficial local flow conditions at angle-of-sideslip. In other cases, such as the F-14 and F-15, the inlets are located on the side of the fuselage. The side-mounted inlets provide shielding at angle-of-sideslip for the windward inlet but not the leeward inlet. In addition, the side-mounted integration may not provide any benefits at high angle-of-attack. The F-18 blends these techniques by utilizing side-mounted inlets shielded by a wing leading edge strake. Other approaches may also be employed to address maneuver effects such as cowl lip shaping or tailoring the engine control system to tolerate maneuvering distortion characteristics Structural design impacts on aerodynamic design Structural integration and manufacturing must also be considered in the inlet aerodynamic design, since aerodynamic surfaces may be constrained by structural design limitations. A good example is provided by Hagseth (1987) for the increased airflow inlet for the F-16. The F-16 was originally designed for the Pratt & Whitney F100 engine, but with the introduction of the General Electric F110 engine and later versions of the F100 engine, the inlet was re-designed to provide increased airflow capacity. The F-16s modular inlet structure facilitated the re-design without impact to the overall airframe structure, but this also imposed constraints on the inlet aerodynamic surfaces. Additionally, inlet structure has been classically composed of machined aluminum bulkheads with formed sheet metal skins requiring labor-intensive installation of thousands of fasteners. Aluminum skins were eventually replaced with composite skins for weight considerations or to improve aerodynamic design flexibility. 2.2 Nozzle & exhaust integration Tactical aircraft exhaust integration is a partnership between propulsion system and air system contractors. The exhaust nozzle is typically designed and manufactured by the PSC, and furnished as engine-mounted hardware. Nevertheless, since exhaust nozzle design and aftbody integration are important factors in propulsion system performance and aerodynamic drag, the air vehicle manufacturer has a large role in specifying the nozzle shape, size, and functionality. It is the air vehicle manufacturer s responsibility to ensure the nozzle and aftbody are integrated with minimum drag Axisymmetric convergent-divergent nozzles Most classic tactical aircraft exhausts systems are axisymmetric for structural efficiency with variable geometry for afterburning. The exhaust nozzle contains a convergentdivergent (CD) internal flow path of overlapping flaps to achieve variable throat and exit areas. Overlapping external flaps provide a fairing between the nozzle exit and the air vehicle aftbody. In addition, the exhaust system for tactical aircraft is typically designed to allow afterburning, which requires a large variation in the minimum nozzle flow area (the throat area, or A 8 ) for operability. In non-afterburning mode (known as dry power ) the throat area is sized to maintain stable, efficient engine operation. In afterburning, the throat area must be substantially increased to compensate for decreased flow density. Nozzle exit area, A 9, is also varied with afterburning to optimize installed performance. In most cases, A 9 and A 8 are mechanically linked and scheduled with power setting according to an A 9 /A 8 ratio Nozzle jet-induced drag Aftbody and exhaust integration are critical to maintaining low drag. There is a trade between nozzle length and aftbody

5 Tactical Aircraft Aerodynamic Integration 5 boattail angles and the resulting drag and weight. Finding the optimum length and external shape is important to net propulsion performance. For example, Catt, Welterlen and Reno (1993) demonstrated that exhaust nozzle shaping and length for the F-16 axisymmetric nozzle are critical to minimizing drag and jet effects, and Schnell (1974) demonstrated aircraft minimum drag reductions through nozzle shaping for the F-14. Additionally, the aftbody is typically not axisymmetric, and integration of an axisymmetric exhaust may result in large aftbody boattail angles and base regions, which introduce drag. This is evident in twin engine configurations such as the F-15 where a base region is required between the two closely spaced nozzles. These base regions can be used as an exit for secondary flow systems, since they typically feature low local pressure. For example, the F-16 base regions on either side of the exhaust are used for nacelle ventilation exits. Other integration approaches include 2-D or more highly integrated exhausts for survivability as with the high aspect ratio F-117 exhaust (a non-afterburning design). Such conformal exhausts can provide a better integration with lower drag or survivability benefits, but they are less efficient structurally, and therefore heavier, than axisymmetric designs. These types of exhausts are typically avoided due to complexity and weight Thrust vectoring The exhaust system can be used to enhance air vehicle maneuverability through the use of thrust vectoring (TV) deflection of the exhaust jet relative to the airframe to produce a non-axial thrust component and a net control moment on the aircraft. TV typically requires mechanical variation of the exhaust system to accomplish either single axis (pitch or yaw) vectoring, or a combination of both pitch and yaw, known as multi-axis thrust vectoring (MATV). Thrust vectoring schemes have been pursued for many years, with numerous concepts proposed, investigated, and brought through various stages of development. Several MATV approaches using axisymmetric nozzles have been demonstrated through flight test maturity, including the F-18 High-Alpha Research Vehicle (Asbury and Capone, 1995) and the F-16 Multi-Axis Thrust Vectoring configuration (Small and Bonnema, 1994). Regardless of the approach, TV has been demonstrated to provide substantial benefits in maneuverability. However, due to weight, complexity, the application-specific desire for close air combat maneuvering, and the potential redundancy of TV with aerodynamic control services, thrust vectoring has not been implemented in production for legacy US tactical systems. The recently developed F-22 proves an exception to this rule, however, as it employs 2-D nozzles with pitch axis vectoring (discussed in more detail in Section 6.1.2) Ejector nozzles Integrated ejector capability is another example of an exhaust nozzle feature for which very close coordination is required between PSC and ASC. Ejector nozzles (in which a secondary flow is pumped by the primary jet flow) are typically used for exhaust system cooling, pumping secondary flow, or airflow matching. For example, the TF30 engine for the F-111 employed an ejector nozzle with air supplied from the freestream for aerodynamic area control. During the 1990s, several efforts were undertaken to demonstrate the viability of an ejector system for cooling the F-16 nozzle to improve nozzle flap durability and thus reduce support cost and logistics footprint. In these efforts, both Pratt & Whitney F100 and General Electric F110 engines were modified to incorporate an ejector slot on the internal flaps slightly downstream of the nozzle throat. Ejector cooling flow was supplied by an external secondary air inlet supplying the nacelle ventilation system. A major challenge with ejector design is matching the supply air pressure to the ejector slot pressure to provide adequate airflow pumping over a wide range of flight conditions. While each F-16 design worked as planned, the benefits of nozzle cooling were not shown to outweigh the added weight and complexity of the system, and it was never adapted for a production version of the aircraft. 3 SYSTEM ENGINEERING REQUIREMENTS Recent practice in the United States has moved to a performance-based specification (PBS) approach in which key performance parameters (KPPs) are specified at the air system level. This approach allows for much greater interpretation by the air system contractor as to how the air vehicle is configured, and how systems are balanced to provide an overall optimization. Systems engineering requirements allocation techniques have thus become critical to correctly decompose KPP requirements to the system, subsystem, and component level. In this section, some of the more important requirements related to tactical aircraft propulsion aerodynamic integration are discussed. Of particular note, survivability and affordability have emerged as leading design drivers. 3.1 Classic aeroperformance: Net installed thrust, TSFC and operability While inlet and nozzle aerodynamic performance metrics such as inlet pressure recovery and nozzle thrust

6 6 Propulsion Integration coefficient remain as important measures of component performance, the concept of net installed propulsive thrust (Fnet) is more useful in judging the overall performance of the installed propulsion system and in allocating performance requirements to lower-level components. Fnet includes consideration of uninstalled engine thrust, free stream momentum (known as ram drag ), inlet pressure recovery, inlet spillage and/or bypass drag, inlet boundary layer control drag, nozzle internal performance, nozzle jet-induced aftbody drag, and the associated secondary air system drag components. Fnet also often accounts for a nominal level of power take-off from the engine, be it in terms of compressor bleed or mechanical horsepower extraction. It must also be noted that the propulsion system integration will have an impact on overall non-throttle dependant aircraft drag, C dmin. This dependency becomes more entwined as design of inlets and nozzles become more integrated; thus, it is critical early in the development process to identify a rigorous aero/propulsion bookkeeping scheme. Likewise, as propulsion and power/thermal management systems become increasingly integrated, similar bookkeeping issues are introduced. Net installed thrust has a first-order impact on air vehicle-level performance metrics such as acceleration time, maximum speed, and specific excess power. In achieving the highest performance levels, the ratio of F net to aircraft weight will exceed 1.0. To optimize Fnet, it is critical to achieve an efficient inlet and nozzle aerodynamic integration, since installation losses can reduce uninstalled engine gross thrust on the order of 30 50% or more. Thrust specific fuel consumption (TSFC), defined as the ratio of fuel flow rate to net installed thrust, is likewise of first-order impact on air vehicle-level performance metrics such as range. Additional discussion on these parameters can be found in Basic Principles: Thrust, Drag and Induced Forces. The installed propulsion system configuration must also maintain engine operability across the entire range of aircraft operation. Inlet distortion (spatial, temperature, planar wave, etc.) and low airflow instabilities are the primary operability concern driven by aerodynamic integration; see Basic Principles: Gas Turbine Compatibility. 3.2 Survivability Incorporation of survivability (also called stealth or Low Observable ) techniques is now recognized as a key requirement for current and future tactical aircraft designs, with consideration of radio frequency, infrared, visible, and acoustic regimes all of some level of emphasis. Survivability considerations have created a dramatic shift in propulsion integration techniques as well as overall air vehicle design. While aeroperformance is still the most dominant requirement for the propulsion system and overall air vehicle, certain fundamental survivability features are evident as state-of-the-art on many modern military platforms. These techniques include edge alignment, sloped/faceted surfaces, engine obscuration, etc. The F-117 Stealth Fighter and B-2 bomber exemplify how the entire air vehicle design can be dominated by survivability considerations, resulting in nontraditional propulsion concepts and air vehicle shaping. Since specific design techniques and related performance are tightly protected by aircraft and engine manufacturers and their sponsoring organizations, a thorough discussion is out of scope of this article. Nevertheless, it can be stated that there is a substantial and unprecedented design challenge incorporating survivability features while retaining high aerodynamic performance and simultaneously improving affordability. 3.3 Weight Weight has always been a parameter of high concern in any tactical aircraft design, since it has first-order impact on both vehicle performance and procurement cost. Inlet weight is somewhat difficult to quantify, since much of the inlet is highly integrated with the overall airframe. It is typical to look at inlet weight on a delta from baseline basis and consider those discrete parts that can be severed (e.g., dedicated aperture or duct structure, mechanical actuation components, ducting, doors, screens, etc.). Inlet weight is driven by physical cross-sectional size (as determined from engine and secondary airflow requirements and the sizing approach), length (which falls out from the overall integration scheme), mechanical complexity, and structural loads. While steady-state operating pressure and aircraft maneuver loads are important, it is the hammershock load that results from an engine stall that is usually the structural sizing case. Tactical aircraft engine exhaust systems are usually provided as part of the engine, and as a result, their weight is bookkept and optimized at the engine system level. However, air vehicle system level requirements, including survivability, functionality (i.e., thrust vectoring or ejector cooling), mold line integration and drag reduction, and mechanical control range and rates, can have a first-order impact on weight. 3.4 Affordability and life cycle cost Life cycle cost (LCC) is defined as the summation of development, acquisition, and operation and support (O&S) costs for a given system. LCC is a critical element in the concept of

7 Tactical Aircraft Aerodynamic Integration 7 affordability, which is loosely defined in terms of a capabilityto-cost relationship, or more simply, as the cost of obtaining a certain combat capability. In recent years, affordability has become a dominant consideration for tactical aircraft, and this requirement has flowed to the propulsion system. As with weight, it is difficult to separate the LCC contribution for a propulsion component or subsystem from that of the system as a whole. Furthermore, while development cost itself may represent only a small portion of the total LCC, both acquisition and O&S costs are to a large extent defined by the propulsion configuration features selected early in development. Acquisition cost is a strong function of physical size, shape, weight, material selection, mechanical complexity, and manufacturing difficulty, all of which are determined in the design phase. O&S cost for inlet or nozzle components is even harder to quantify, but in general trends with mechanical complexity and reliability, vulnerability and damage tolerance (including foreign object damage, FOD), material reparability, etc. Since fuel cost is part of support cost, fuel consumption characteristics of the propulsion system also influence LCC. As an example trade, a variable-geometry inlet with mechanical actuation might provide better aerodynamic performance leading to better TSFC and lower fuel cost, but at the same time, it will weigh more, cost more to build, and require more maintenance events and hours per event to keep all its moving parts in repair. 3.5 Noise, emissions and operational/environmental robustness Excluding stealth considerations, propulsion noise (jet noise and engine/fan noise) and CO 2 and NOx emissions have not been of high concern for tactical aircraft as other parameters described herein. Few compromises to either legacy or modern air vehicle systems were made with these parameters in mind. However, as civilian communities become more involved regarding peacetime basing options for fighter aircraft, noise is becoming more important. Also, as a result of both rising price of fuel (and thus O&S cost) and the Green movement to reduce aircraft emissions, fuel efficiency is becoming of greater interest to military customers. Inlet and nozzle systems must also be robust against a host of operational/environmental considerations. Such considerations include bird strike; ice accretion, shedding, and ingestion; salt water, aircraft fluids, and other corrosive substances/chemical agents; electro-magnetic interference (as applied to electronic control systems); armament exhaust or other hot gas ingestion; foreign object debris including sand/particulate matter ingestion and resultant erosion/fouling potential; extremes of temperature (especially the impact on actuation fluid systems); maintenance, repair, and overhaul actions (maintainer access, dropped tool impacts, seal degradation); structural loads; vibration/acoustic loads (especially external turbulence), etc. Design response to these requirements is typically based on legacy techniques and best practice. 4 ADVANCED INTEGRATION TECHNOLOGIES During the last 20 years many advanced aerodynamic integration technologies for tactical aircraft application have been identified, researched, developed, and demonstrated. Throughout the 1980s, the primary objective was improvements in aerodynamic performance and/or functionality (such as thrust vectoring) and stealth capability. As the Cold War ended and defense acquisition budgets were reduced, the emphasis shifted to decreased weight and life cycle cost while maintaining aerodynamic performance and stealth capability. While many viable advanced technologies have been matured over the past two decades, there have been limited opportunities to integrate those technologies into production systems due to the limited number of new or derivative tactical aircraft programs. Although an advanced technology may clearly improve air vehicle capability and LCC, it also introduces additional development risk and cost. Consequently, advanced technologies typically require substantial demonstration and evaluation to reduce risk prior to introduction into a production program. Flight demonstration is typically viewed as the culmination of a technology development program (Moorehouse and Hamstra, 2003). Even with adequate risk reduction, any given technology must justify inclusion into a production system based on merit shown through system-level trades on that specific application. Some of the technologies discussed in this section have been incorporated on production tactical systems; others may demonstrate benefits for future systems, may have utility for strategic, mobility, reconnaissance or other systems, or may simply remain on the shelf. 4.1 Inlet technologies Advanced inlet shaping/integration concepts such as the caret inlet and diverterless inlet are excellent examples of approaches to achieve high aeroperformance with shaping for survivability. Other inlet technologies of particular note include compact, full-obscuration inlet ducts, inlet flow control, and structurally integrated inlets.

8 8 Propulsion Integration Caret inlet The caret inlet technique has been understood as an academic concept for many years (Seddon and Goldsmith, 1985), but was not matured to a realistic engineering design until the 1980s. The primary trait of caret inlets is a pair of oblique compression ramps that generate a 2-D flow field and coplanar shock waves at the supersonic design point. Primary advantages of the caret inlet are efficient supersonic flow compression (as with the F-14 or F-15) and swept inlet edges that can be aligned with the aircraft planform. The challenge with the caret inlet lies at supersonic, off-design conditions where the shocks generated by the two ramps are no longer co-planar, resulting in shear layers and potential distortion and inlet instability. The caret inlet concept was adopted for both F-22 and F/A-18E/F. Design and development of the F/A-18E/F inlet is discussed in Hall et al. (1993) Diverterless inlet Another advanced inlet integration approach investigated and developed during the 1990s was the bump inlet. This approach was of interest due to its potential to allow elimination of boundary layer diverters. Whereas the caret inlet is based on a 2-D flow field, a bump inlet is derived by streamline tracing through a three-dimensional flow field (Seddon and Goldsmith, 1985). Boundary layer diverters have traditionally been crucial for most tactical aircraft inlet integration approaches. The diverter serves to improve performance and maintain supersonic inlet/engine compatibility by preventing boundary layer ingestion and physically isolating the inlet shocks from the forebody boundary layer. However, diverters may introduce undesirable survivability characteristics and inhibit physical integration of the inlet cowl and forebody, which is desirable from a weight reduction standpoint. Various research efforts were undertaken to mature the bump inlet concept into a practical design, including a flight test effort on an F-16 that lead to incorporation of such an inlet on the JSF X-35 concept demonstrator aircraft and production F-35 aircraft (see Hamstra, McCallum and McFarlan, 2003; and Hehs, 2000). This particular design, known as the diverterless supersonic inlet, integrated a highly three-dimensional bump compression surface with a forward-swept cowl. This combination produces a pressure gradient that diverts the majority of the boundary layer and provides a stable interaction between the inlet shocks and remaining boundary layer, eliminating the need for both boundary layer diverter and bleed systems. A diverterless inlet concept was also employed on the JSF X-32 demonstrator Compact inlet ducts As tactical aircraft design evolved in the 1990s, compact inlet ducts became another technology of emerging interest due to the desire to enable lower cost through lower inlet length. In this context, compact refers to short inlet ducts (L/D 4) that achieve full line-of-sight obscuration of the engine, which is necessary for survivability compliance. Unfortunately, achieving full obscuration in a compact design requires high rates of duct curvature and flow area/shape change, all of which traditionally introduce unacceptable pressure loss and distortion. However, through the use of modern design techniques, researchers were able to develop and validate with wind tunnel experiments compact designs that achieved excellent performance (see Philhower, Robinson and Brown, 1998) Inlet flow control In the 1990s flow control began to emerge as a technology with many aircraft applications. In the context of jet engine inlet systems, flow control refers to the manipulation of large-scale flow phenomena (such as inlet distortion) with relatively small-scale perturbations to the flow enacted in highly-receptive zones within the inlet. Flow control techniques may be active or passive and involve open loop or closed loop controls. One example of inlet flow control is for reducing adverse secondary flows in ultra-compact inlet ducts (full obscuration designs of L/D<4), suggested by Anderson et al. (1999) and continued by Miller, Anderson and others (see Hamstra, Miller and Truax, 2000). These efforts showed that as duct curvature increases, the secondary flow increases to a point where a vortex will lift off the duct surface, thereby decreasing pressure recovery and increasing distortion at the engine face. Studies found that localized placement of microvanes or airjets could control the secondary flow and prevent vortex lift off. Inlet flow control research continues and has matured through full-scale demonstration of a representative advanced inlet coupled to a turbofan engine Structurally-integrated inlets and probabilistic loads Structurally integrated inlets, including unitized, fastenerless structure based on probabilistic design loads, were also developed to reduce weight and LCC. As mentioned previously, one of the primary design loads for inlet structure is the pressure load (hammershock) due to an engine stall. For legacy fighters such as the F-16, structural sizing was based on a worst case stack up of conditions that would

9 Tactical Aircraft Aerodynamic Integration 9 give the highest possible hammershock load. The worst-case hammershock loads occur at high dynamic pressures, supersonic speeds, and low altitude. However, the vehicle residence time and probability of a stall at these conditions is very low. For more modern designs, research has shown that a probabilistic hammershock load is more appropriate (Gridley, Sylvester and Truax, 1999). For the case studied and based on an acceptable risk of occurrence of 1 in 10 million, this approach reduced design loads from 70 to 44 psid, which resulted in an estimated inlet duct weight reduction of 40%. Inlet structural design and manufacturing technologies similarly evolved to consider structurally integrated concepts to further reduce weight and LCC. Modular legacy inlet system components, which were bolted to the airframe or assembled as part of a fuselage section, were replaced with unitized components sharing airframe bulkheads and loads. As mentioned in Section 2.1.6, inlet structure for legacy fighters consisted predominantly of aluminum bulkheads and frames and multi-piece aluminum or composite duct skins attached with thousands of fasteners. In the most advanced approaches considered, the multi-piece aluminum and composite skins are replaced with single-piece, fiber-placed composite ducts. Metal inlet duct bulkheads and frames are replaced with composite webs with preformed, bonded joints, which reduce/eliminate fasteners and integrate directly with airframe bulkheads. This approach reduced cost substantially by eliminating the labor-intensive fastener installation. However, these concepts required expensive tooling for the single-piece duct, and repairs could become problematic in the event of extensive damage to inlet structure. 4.2 Nozzle technologies Since the early 1980s, R&D efforts continued to improve the classic convergent-divergent (CD) axisymmetric nozzle in terms of affordability, survivability, and aftbody drag reduction. Engine manufactures have made enough incremental progress in these areas such that the engine-mounted axisymmetric CD nozzle remains as the dominant architecture for the most modern tactical aircraft. At the same time, however, structurally integrated, conformal exhaust systems were investigated to increase survivability and reduce weight. Conformal shaping in this context refers to exhaust systems that are highly integrated with the natural shape of the vehicle planform, such as those on the F-117 and B-2. Conformal shaping and partially or fully fixed internal and external geometry simplify aftbody integration and allow elimination of gaps, seals, and visible moving parts, but introduce enormous challenges regarding throat area and thrust vector control. As an ultimate goal, nozzle technologists have desired to combine the shaping features of F-117/B-2 nozzles with the vectoring/afterburning functionality of the F-22 but to do so at the cost and weight of a traditional axisymmetric nozzle. This goal remains elusive Advanced vectoring techniques Requirements for thrust vectoring complicated the task of developing fixed-geometry exhausts. While pitch/yaw vectoring was investigated, the primary approach was yaw vectoring for control of tailless aircraft. Two successful approaches emerged from multiple investigations: the nozzle throat disc and fluidic injection (or nozzle flow control, see Section 4.2.3). The nozzle throat disc (Garret and Zilz, 1992) employed a rotating disc at the throat station to skew the sonic line, thus creating vectored thrust. This approach achieved high levels of vectoring, but, due to the variable geometry throat, weight was still an issue. The fluidic vectoring approach (Miller et al., 2000; Miller, Yagle and Hamstra, 1999) employed fluidic injectors (using engine bleed) in the nozzle expansion section to either skew the sonic line or create a weak oblique shock and thus vectoring, while simultaneously controlling effective jet throat area. This approach achieved high levels of vectoring and is lighter, since it is compatible with a physically fixed geometry throat. However, fluidic approaches require a high-pressure airflow source that must be powered and controlled, thereby introducing a new set of issues. Many other mechanical and fluidic approaches remain under investigation Exhaust structural integration Structurally integrated exhaust systems have been investigated to reduce weight and LCC through load sharing between airframe and exhaust structure. Such load sharing is challenging due to non-uniform thermal expansion and high thermal loads. Various approaches, including engine-mounted, airframe-mounted, and structurally integrated nozzles, have been utilized for advanced aircraft with varying degrees of success. Despite significant advancement in state-of-the-art, structural integration of hot exhaust structure with airframe structure remains a very difficult problem Nozzle flow control Active flow control has been an area of significant investigation, with the desire to provide both throat area (A 8 ) and/or thrust vector control (see Section 4.2.1). Flow control has also been shown as applicable to aftbody drag reduction (see Haid and Gamble, 2004) and jet noise reduction. For nozzle as

10 10 Propulsion Integration well as inlet systems, a major issue is the penalty introduced for extracting mechanical, pneumatic, or electrical power to drive the flow control technique. As mentioned in Section 3.1, power extraction results in an adverse impact on net installed propulsive force, and must therefore be justified at the vehicle system level. Active flow control may also involve routing, valving, insulation, software controls, and failure mode considerations, all of which must be factored into affordability and survivability improvements. 5 EVOLVING DEVELOPMENT TECHNIQUES In the time period between legacy and modern tactical aircraft, numerous advancements have been made in the design engineer s ability to produce, analyze, and harness information describing the physical and performance characteristics of inlet and nozzle systems. Most of these improvements resulted from advancements in information technology powered by gains in computer processing hardware and the associated software and infrastructure. Four specific areas of advancement were computer-aided design (CAD), computational fluid dynamics (CFD), computer-aided manufacturing (CAM) and rapid prototyping, and high-speed experimental data acquisition and analysis. A detailed perspective on these techniques can be found in Raj (1998). Perhaps the most drastic change regarding inlet and nozzle design involves the use of CFD methods (see CFD). While CFD has advanced in terms of solver accuracy, user interface, hardware cost, graphical output, and so on, it has been solution speed (e.g., turn-around time) that has the most significant impact. Speed in this context implies more than performance indices obtained in a laboratory environment. Rather, speed is measured by the solution throughput achievable on a day-to-day basis by design engineers with access to typical, non-exotic computing resources for example, the number of solutions of a given size that can be converged overnight on an individual engineer s workstation. Various studies have shown that this trend follows a Moore s Law - like exponential improvement with time. With improvements in solution speed and results analysis capability, the aerodynamic designer is able to derive increasingly sophisticated insights into the flow field phenomena and harness that knowledge for design improvements. 5.3 Computer-aided manufacturing CAM techniques, in particular, numerically controlled (NC) machining, have contributed to development of advanced inlet/nozzle integration schemes by providing improvements in wind tunnel model mold line accuracy, parts interchangeability, rendition of sophisticated contours, and so on. In the 1990s, rapid prototyping techniques, such as laser stereo lithography and direct manufacturing, also became popular as a low-cost, rapid turn-time method for manufacturing wind tunnel model parts or even complete models. 5.1 Computer-aided design Three-dimensional CAD tools, often coupled with structural or aerodynamic analysis tools, have provided the ability to craft complex geometries with little effort, and to store and redesign those items with ease. As an example, modern CAD tools allow lofting of engine inlet ducts in 3-D space following specified centerline, flow area, and cross-sectional shape distributions. The duct contour can become very sophisticated when short, compact full-obscuration designs are required. Some of these designs would be difficult if not impossible to loft without computer methods. Additionally, parametric models allow the designer to quickly explore and modify designs in minutes instead of hours or days. 5.2 Computational fluid dynamics 6 MODERN TACTICAL AIRCRAFT CASE STUDIES: F-22 AND F-35 The two most recent examples of clean-sheet tactical aircraft development in the United States, the F-22 and F-35, are discussed as state-of-the-art case studies. The F/A-18E/F, an evolutionary upgrade to the F/A-18C/D, is also noteworthy and is discussed in detail by Hall et al. (1993). One can see that with the F-22 and F-35 inlet and nozzle designs, top-level systems engineering requirements such as stealth compatibility, weight, and procurement and support cost had a major impact on the overall system design. 6.1 F-22 Advanced tactical fighter The F-22 is a twin-engine air superiority aircraft featuring a classic wing/body/tail configuration shaped with stealth requirements in mind. A prototype YF-22 first flew in 1990 and the first production version flew in The F-22 was

11 Tactical Aircraft Aerodynamic Integration 11 Figure 3. F-22 Propulsion aerodynamic integration features. designed to replace the F-15 as the front-line fighter for the United States Air Force. The F-22 incorporates twin Pratt & Whitney lb (160 kn) thrust class F119-PW-100 twin spool afterburning turbofan engines. With these engines, the F-22 is able to sustain supersonic cruise speeds without the use of its afterburners. Additional details on the evolution of the F-22 aircraft and F119 engine can be found in Hehs (1998) and Deskin and Yankel (2002), respectively. Key F-22 propulsion integration features are shown in Figure The F-22 engine inlet The F-22 inlet system features twin side-mounted inlet apertures, each feeding a single engine through a long (L/D 6), full-obscuration S-duct. The apertures are highly swept for stealth compatibility and utilize a two-dimensional, external compression Caret design (see Section 4.1.1). Boundary layer control is provided with a classic boundary layer diverter and bleed system. The inlet design provides an inherent angle-of-attack shielding function, enabling full engine/inlet compatibility across a wide range of operating conditions. Due to stealth, weight, and mechanical reliability (i.e., maintenance cost) considerations, the F-22 inlet geometry is fully fixed, and does not have the variable geometry features of previous air superiority systems such as the F-14 and F-15. F-22 inlet development was aided by significant advancements in wind tunnel test data processing, specifically, the ability to produce near-real-time analysis of dynamic distortion data. Prior to the F-22 timeframe, dynamic distortion data was recorded during the test but processed and analyzed post test. With improvements in computer power, by the late 1980s hardware was available to acquire and analyze dynamic data at the test site, allowing engineers the ability to make design decisions based on this data in near-real time. While CFD analysis was used on F-22, solution throughput in the late 1980s/early 1990s had not yet reached the rate required to base major decisions on CFD-produced information (in the absence of supporting test data). The role of CFD was to add information to a knowledge base derived primarily from test data The F-22 engine nozzle Supplied as part of the twin F119 engines, the F-22 nozzle features a stealth-compliant, 2-D, convergent-divergent, thrust vectoring design. The nozzles are highly tailored to the F-22 s requirement for optimum performance (net installed propulsive thrust (Fnet)) at minimum weight. In-flight thrust vectoring enables an enormous Mach/angle-or-attack/angleof-sideslip operating range, which in turn creates a maneuver compatibility requirement for the engine inlet system. With the mechanism in place for thrust vectoring, added capability for thrust reversing was also studied early in the program, but this feature did not trade favorably with cost and weight impacts, and thus is not present on the production system. 6.2 F-35 Joint strike fighter The F-35 is a single-engine tactical aircraft featuring a wing-body-tail configuration layout similar to the F-22. A demonstrator X-35A version flew in October 2000 and the first production version flew in From inception, the aircraft was designed with three specific variants in mind to meet the needs of US and allied Air Force, Navy, and Marine customers Conventional Takeoff and Landing (CTOL),

12 12 Propulsion Integration Figure 4. F-35 Propulsion aerodynamic integration features. Short Takeoff Vertical/Landing (STOVL), and aircraft carrier variant (CV). Each variant is designed to unique requirements, yet features high commonality with the other variants to enable overall affordability. The F-35 is thus designed to replace the F-16, F-18, and AV-8B and numerous other aircraft. Two interchangeable lb (180 kn) thrust class afterburning turbofan engines can power the F-35, either the F135 offered by Pratt & Whitney or the F136 offered by the General Electric/Rolls-Royce Fighter Engine Team. CTOL and CV propulsion systems are identical; the STOVL variant incorporates a shaft-driven lift fan in the forward fuselage, a vectoring engine nozzle, and other special systems to provide thrust augmentation and aircraft control necessary for STOVL operation. Key F-35 propulsion integration features are shown in Figure The F-35 engine inlet F-35 engine inlet design considerations included stealth compatibility, cost and weight, transonic cruise and maneuver performance, and amenability to low-speed performance enhancement for STOVL operation. The engine inlet system features twin side-mounted external compression inlet apertures and a bifurcated inlet duct. All three variants have identical engine inlet systems except for incorporation of a top-mounted auxiliary inlet on the STOVL aircraft. A twin side-mounted inlet arrangement was chosen for compatibility with the overall aircraft configuration, in particular, the centerline-mounted STOVL lift fan. The inlet aperture utilizes a highly 3-D compression surface and forward-swept cowl to provide both flow compression and boundary layer diversion functionality. These features allow elimination of a discrete boundary layer diverter and/or boundary layer bleed system, thereby improving stealth compatibility and reducing inlet weight, manufacturing complexity, and cost while maintaining required performance (total pressure recovery and drag) and compatibility (distortion and airflow matching) levels (see Section 4.1.2). The primary engine inlet is fully fixed with no moving parts. A very short bifurcated S-duct fully obscures direct view of the engine. Advanced CFD techniques played a critical role in development of the F-35 inlet system. The highly contoured design of the inlet compression system cannot be assessed by conventional analytical techniques. Similarly, the compact inlet S-duct features complex flow characteristics that are sensitive to subtle changes in duct area, offset, shape, and inflow distortion pattern, and thus require computational means to model. Advancement in wind tunnel test techniques was also required by the F-35. With both lift fan and main engines supplied by closely coupled inlets, a standard 40- point instrumentation rake was shown to be inadequate to render distortion patterns to necessary resolution, and new instrumentation designs thus had to be developed The F-35 engine nozzle Key F-35 engine nozzle design considerations included stealth compatibility, cost and weight, STOVL compatibility, and traditional nozzle performance. As part of the aircraft s engine interchangeability requirements, both F135 and F136