Finally, LM's MQ-XX design candidate is revealed; BWB type wing-fuselage design with one engine can be noticed. Asymmetric layout of aerial re-fuel tank can be specified.
At the nose, several sensors are located for (possibly) navigation, simple surveillance, or situation awareness while some antennas are located under the fuselage, and fuel tank is located beside of it.
Landing gears are well designed to secure space for additional payload on the wing, however, attachment of additional payload is uncertain. Inlet duct are not severely distorted, usual for stealth aircraft with loss of engine efficiency, to obtain better cruise efficiency with small loss of duct.
Compare to Boeing's design, it looks lack of canted VT and huge fuselage. It looks similar to the company's previous RQ-170 while Boeing made it looks like Tacit Blue (Although Tacit Blue was from NG)
As a candidate of surveillance aircraft of USN, Boeing had proposed joint wing design. Most of the sensors are mounted on the wing, and Boeing insists it could cover and provide better sensor capability than existing systems.
It should be noted that complex flow pattern at the wing tip of the joint wing is studied via wind-tunnel testing. Fuselage design is similar to S-3, and engine pod is attached at the side of the body where downstream of the wing is worried.
Joint wing is expected to have advantage in lift/drag ratio and structural fatigue aspect, and this is why Boeing chosen this design for long-time endurance aircraft. But, sill, complex control of CG, and engine position could be trouble some phenomena when the design of the aircraft is matured.
Additionally, rear wing part tends to be lower or higher position than front wing to avoid downstream of the front wing which degrades performance and handling qualities or stability. In that design, if rear wing is swallowed by downstream of the front wing, aircraft could enter the severe departure status than the conventional ones.
Before start of Luna's New year in Korea, Beoing revealed its MQ-XX candidate prototype. Although it has stealthy outer line, Boeing denied its possibility of stealth performance. It has high AR main wing and v-tail which focused its performance on the long-endurance 'tanker' mission. Maybe Northrop Grumman and Boeing compete for MQ-XX of USN; Avenger C can also be candidate from General Atomics. This is the start of the 'unmannization' of USN; MQ-XX probably takes role of tanker, submarine-tracking, light-ISR depending on its mission equipment. Next step of this development leads to take role in delivering of simple weapons such as cruise missiles or AG weapons. Frankly, X-47B shape has more potential for development of LO performance aircraft required for 'future-missions'. Boeing's one is more optimized for tanker mission rather than other risky missions. Still, there is no certain which company takes winner position of this competition. Northrop-Grumman has experience in test of unmanned vehicle in USN, however, no-one easily expect winner.
Recent release from the Russia shows hypersonic weapon development, air-launched cruise missile from MiG-31. Class of missile is hypersonic cruise missile targeted aircraft carrier group; penetrating Aegis system of USN. Missile looks like size of Kh-41, and having network? or avoidance technology same as conventional Russian anti-ship missiles. It is interesting that how much Russian matured scramjet technology. (and communication technology at hypersonic speed)
New concept of cruise missile targeted replacing famous Exocet missile proposed in 2011 Paris airshow. Supersonic air-breath engine (seemed like ramjet type) is attached beneath the fuselage.
Interesting thing of this missile is that two additional ammunition is released from the side bay of the missile. This sub-missile is guided by inertial guide system. Purpose of the two sub-missiles in suspicious.
PART C. Stability / Quality and Structural Issue via High AoA flow
Other than simple criteria or analytic methods, direct observation by wind-tunnel, CFD, model-flight, experimental flight of full scale aircraft could be performed. These experiments are also taken for calculation of static coefficient used by criteria and analytic methods, however, recently, enhanced methods are provided to give direct insight or full 6DOF motion of jet fighter departure. This session introduces you to part of the references which attempt to analyze high AoA motion other than simple criteria or analytic methods.
Most common method is wind-tunnel testing for static or dynamic model [3-13]. Reference  showed one of the recent measurement method for the wind-tunnel model which uses pressure sensitive paint (PSP) on the model. Left part of the Fig. 2.16 presents ‘frequency gate’ for absorption and illumination band frequency; the paint absorbs specified frequency of light for certain time, then it illuminates the light. The illuminated frequency of the light is changed by the applied aerodynamic pressure captured by high resolution camera as shown in the right of the Fig. 2.17. This method has advantage that can track pressure change of local area of aircraft. It means that complex phenomenon in high AoA such as vortex breakdown, adverse gradient of the pressure, separation could be directly or indirectly shown as Fig. 2.18. Fig. 2.18 is a typical example of the result of the PSP experiment for Alpha-jet aircraft which represent low pressure peak on the wing and high pressure concentration on the wind-shield. Pressure variation could reveal not only high AoA flow change also shock wave via Mach number effect.
Indeed, PSP method helps to show details of the flow around the jet fighter, and correction for the sting effect is possible by using the method. However, as shown in Fig 2.17, many high resolution camera and illumination devices are required to capture the image; naturally these devices need high cost and post-process devices to analyze image. When the attitude of the aircraft, AoA and side-slip angle, is changed, relative angle for the camera is also changed and could affect the quality of the data, effect of image distortion. Because of this characteristics, lumped data, lift, drag, and moment coefficient, are usually tracked by force sensor located inside of the model and sting.
Fig. 2.16 Introduction of pressure sensitive paint (PSP) method for wind-tunnel experiment 
Fig. 2.17 Typical PSP experiment result for Alpha-jet, developed by Franco-German; it showed pressure change via lift and drag
Free-to-roll sting method is also famous for F-35 development  for wing-drop problem analysis. In that method, model is installed on the sting, and model is free for the roll rotation direction of the sting as shown in Fig 2.18. It has strong point in evaluation of transverse direction stability. As shown in Fig. 2.19, various options had been tested for F-35C model including array of vortex generator, sealing of gap between outer-inner leading edge flap. Similar test was already taken for F/A-18E  (not free-to-roll) as shown in Fig. 2.20; effect of complex flap configuration of Super-Hornet was tested for maneuver. Conceptual studies using wind-tunnel testing have been performed for leading edge flap effect ; interaction between LERX and segmented leading edge flap for lift and drag of the aircraft (Fig 2.21). Change of this design could affect the vortex distribution (LERX) and local AoA on the wing (segmented leading edge flap). Detailed result of these design change will be discussed in Chap 2.2 (also for later references).
Similar study for Eurofighter configuration  focuses impact of design changes on high AoA stability criteria, and it resulted in control surface layout of Eurofighter (Fig. 2.22). Vortex interaction between canard and main wing was also evaluated in the past study for AJS-37 Viggen design ; it resulted bizarre wing shape of Viggen. Cut-off shape of inner part of main-wing and canard flap are used to optimize lift/drag and stability characteristics. Modification of F-15 was attempted to increase maneuverability; micro-blowing , canard and additional strake  as shown in Fig. 2.23 and 24. These tests are conducted in wind-tunnel with dynamic sting because high AoA characteristic evaluation.
Fig. 2.18 Schematic of free-to-roll wind-tunnel experiment; DOF of roll direction is free for the model 
Fig. 2.19 Part result of free-to-roll wind-tunnel experiment; effect of sealing gap and vortex generator array is evaluated for un-expected wing-drop problem 
Fig. 2.20 Wind-tunnel experiment configuration for F/A-18E with flap configuration 
Fig. 2.21 Effect of segmented leading edge flap and LERX is studied at wind-tunnel testing 
Fig. 2.22 Effect of segmented leading edge flap and canard is studied at wind-tunnel testing 
Fig. 2.23 Effect of micro-blowing on high AoA stability of F-15 model in wind-tunnel testing 
Fig. 2.24 Effect of additional canard on the high AoA characteristics of the F-15 
Small scaled free flight models are used to evaluate in free condition; cases, effectiveness, limitation are well summarized in reference . These tests are specialized for high AoA characteristic evaluation. Fig. 2.25 shows that model is not chained by its sting which regulate motion or interfere the flow field around the aircraft model. The bottom of the tunnel blows wind which provides free-stream for the model, then model is floating center of the tunnel section. Depending on what constructor does in the experiment, initial rotation components or AoA is given for the model to observe the AoA stability characteristics of the model. In order to evaluate the impact of the control surface deflection, the model with deflected control surfaces is also tested.
In that tunnel, constructors do not only evaluate AoA characteristics, and also tested how the model recover from the unstable AoA status. If the model could not recover from the departure status, flight region of the aircraft should be limited or design of the aircraft should be modified to meet the requirements. Inertia of that kind of model is delicately synchronized to represent dynamic characteristics of the jet fighter as shown in Fig. 2.26. If the size of the model aircraft gets smaller (highly scaled aircraft), distortion via scale effect goes larger, but scaled aircraft cannot be big enough to meet every aspect of the physics because of size limit of the tunnel section.
Important distortion via scale is well represented in Reynolds number effect, unfortunately critical in stability issues. Fig. 2.27 dramatically shows how scaled model could not represent the full scale characteristics. As shown in the previous example (Fig. 2.8 in Part B), Reynolds number component become important when flow separation occurs. Then, scaled model usually fall into stall (full flow separation from the wing and fuselage) in smaller AoA than full scale aircraft. In some special cases, model larger than tunnel section is tested as free-flight remote-control model as shown in Fig. 2.28; this approach mitigates problem of correction of sting effect or size limit via tunnel section. (sting effect correction will be discussed in future article related to minimum drag)
Fig. 2.25 Free testing wind-tunnel model 
Fig. 2.26 Dimensionless number for the scaled model of the aircraft to represent flight characteristics minimizing 
Fig. 2.27 Lift coefficient change via scaled model test; because flow separation characteristics is sensitive to Reynolds number, maximum lift coefficient and flow-separation-occurring-AoA is changed in scaled model test.
Fig. 2.28 Free-flight remote control model aircraft for F-15, X-31, F/A-18E/F;
Recent development of CFD technique also help to investigate characteristics of the high AoA around the jet fighter [14-19]. Wooden et al.  applied CFD on the F-35; they compare Cp distribution on the wing airfoil for various turbulence model as shown in Fig. 2.29. The result showed low pressure peak or lower surface result relatively better fit to experiment result than trailing edge of the upper surface where the developed vortex is break down. Result at trailing edge is very sensitive to constant of the turbulence model which is hard to predict; generation, breakdown, and merge of the vortex on the wing are changed dramatically by setting of the turbulence model.
Dean, J. P et al.  provided full dynamic behavior CFD of the jet fighter during the AoA turn maneuver, and they evaluate lift, drag, and moment coefficient for the experiment results. Lift and drag coefficient result is relatively well fitted to the experiment data in ±10% error range however, pitching momentum is not easy to fitted as shown in Fig. 2.30. Jeans, T. J et al.  analyzed combination effect of the chinned fuselage and the main wing of the generic jet fighter shape. Delayed detached eddy simulation (DDES) is used to calculate vortex precisely which combine the advantage of RANS and LES. The conventional RANS solve most of the area, then LES analyze near boundary area to give precise result for the generated vortex. Due to this effort and high resolution grid, CFD result roughly fitted to the experiment one, however, still discrepancy exists for more than 25 deg where usually vortex breaks as shown in Fig. 2.31. Those above results showed recent advance definitely well predict complex vortex around the jet fighter, however still, it requires improvement for abrupt vortex behavior prediction and consequence of the vortex interaction to predict precise momentum value.
Fig. 2.29 Comparison of various turbulence model coefficient performance on CFD; generation, breakdown, and merge of the vortex is very sensitive to turbulence model set up .
Fig. 2.30 Lockheed performed full CFD analysis for F-22 high AoA maneuver; lift, drag, and pitching moment coefficient result is compared to that of the experiment .
Fig. 2.31 Lift, drag, and pitching moment result of CFD for chinned generic fighter 
In the previous paragraphs including Part A, B, and C, I summarized aerodynamic, stability, flying quality issues in high AoA, and introduced methods and recent development to evaluate these issues. High AoA flow problem is not limited in air-flow related ones around the jet fighter. Structural and propulsion issues should also be considered to complete the jet fighter design. Unfortunately, I am not a very expert on structural design issues, and I just could brief introduction for these ones.
High AoA flow on the structure of the jet fighter is very high loading on the jet fighter; it results high lift, drag, and pitching coefficient in same free-stream condition. However, among them, the most critical load on the structure related to the high AoA is dynamic phenomena via vortex. Like equations in Fig. 2.14, aerodynamic load on the wing excites natural mode of the thin structures like main wings, horizontal, vertical tails. In most flight region of the jet fighter, structures are much stubborn than aerodynamic and inertia load; it means natural frequency modes of the structures are higher than that of the loads. However, if vortex with high frequency or aerodynamic load in high speed condition at certain AoA could be matched to the modes, structures could suffer oscillation leading fatigue or failure. It means that long and thin wing like glider have more possibility to fall into that range as shown in Vid. 2.1.
In famous F/A-18 design, vertical tail is swallowed by downstream of the vortex generated by LERX; it is intended to obtain control authority in high AoA. Side effect of this design is that vortex could give fatigue load on vertical tails, then designer decided to add structural stiffener and small vertical strake to scatter the vortex as shown in Fig. 2.32. Scattered vortex break down to several small vortexes and have different frequency which minimize excitation via vortex.
Effect of high AoA is also sensitive issue for the inlet design. Whiford  well summarized effect of inlet on the jet fighter design as shown in Fig. 2.33. Inlet requires elimination of the boundary layer generated by forward fuselage or nose to prevent surge or blade stall problem. Left of the Fig. 2.33 showed remaining flow of boundary layer could affect upper surface of the F/A-18 prototype where vortex is generated. In modern jet fighter design, intake like F/A-18 or Eurofighter type is used because turning flow via nose re-direct flow into air intake naturally. As shown in the center of the Fig. 2.33, local AoA on the inlet is affected by position of the inlet, and side intake faces higher local AoA than the others due to the upstream of the wing. When the throttle set up is changed by pilot’s input, required mass-flow for the inlet is also changed. Changed amount of the flow affect the flow around the inlet, and it should be considered to complete design.
Vid. 2.1 Visualization of aerodynamic flutter
Fig. 2.32 Small vertical strake on LERX of F/A-18 (Left), and structural stiffener of F/A-18’s vertical tails (Right)
Fig. 2.33 Aerodynamic design and inlet interact each other for their performances; especially 
In Chap 1, I have discussed reason for why high AoA is need to modern jet fighters, and I reviewed physics and method to investigate high AoA flow in Chap 2.1. In the remaining Chap 2.2 and 3, I will discuss about solutions for high AoA flow and future insight of high AoA maneuver for the jet fighters. Then, I will finish high AoA Aerodynamics topic.
After this high AoA article, next articles will discuss about minimum drag and SCRAM jet topics.
 Mason, W. H., Some High Alpha and Handling Qualities Aerodynamics, Configuration Aerodynamics Class, HiAlphaBasicsPres
 Nolan II, R. C., Wing-Rock Prediction Method for a High Performance Fighter Aircraft, Thesis of Air Force Institute of Technology, AD-A256 438
 Gustafsson, et al., 2012, External Aerodynamic Pressure Loads for a Rigid Fighter Aircraft, A Comparison between PSP and CFD, ICAS 2012, 28th International Congress of the Aeronautical Sciences
 Owens, et al., 2006, Transonic Free-to-Roll Analysis of the F-35 (Joint Strike Fighter) Aircraft, J. Aircraft
 Banks, D. W., et al., 1997, The F/A-18 High Angle of Attack Ground to Flight Correlation: Lessons Learned, NASA Technical Memorandum 4783
 Henderson, W. P., 1978, Effects of Wing Leading Edge Flap Deflactions on Subsonic Longitudinal Aerodynamic Characteristics of a Wing-Fuselage Configuration with a 44 deg Swept Wing, NASA Technical Paper 1351
 Karling, K., 1986, Aerodynamics of Aircraft 37 – Part 1 : General Characteristics at Low Speed, NASA TM-88403
 Chamber, J.R., 2010, Modeling Flight : The Role of Dynamically Scaled Free Flight Models in Support of NASA’s Aerospace Programs
 Chambers, J. R., 1986, High Angle of Attack Aerodynamics : Lessons Learned, AIAA 86, AIAA 4th Applied Aerodynamics Conference
 Grafon, S. B., et al., 1983, High Angle of Attack Stability Characteristics of a Three Surface Fighter Configuration, NASA TM 84584
 Kraus, W., 1980, Delta Canard Configuration at High Angle of Attack, ICAS80-13.1, 12th ICAS Congress
 Schuster, D. M., and Bryd, J. E., 2005, Transonic Unsteady Flow Characteristics of the F/A-18E Near Abrupt Wing Stall, Flow Induced Unsteady Loads and the Impaction on Military Applications Meeting Proceeding RTO MP AVT 123
 Roos, F. W., 2001, Microblowing for High Angle of Attack Vortex Flow Control on a Fighter Aircraft, J. Aircraft
 Wooden, P. A., et al., CFD Prediction of Wing Pressure Distributions on the F-35 at Angles of Attack for Transonic Maneuvers, 25th AIAA Applied Aerodynamics Conference
 Dean J. P., et al., High Resolution CFD Simulation of Maneuvering Aircraft Using the CREATE-AV / Kestrel Solver, AIAA 2011-1109, 49th AIAA Aerospace Science Meeting, AIAA-2011-1109
 Smith, B. R., 2015, Challenges to the use of CFD in the Military Aircraft Industry, AIAA Sci-Tech 2015
 Jeans, T. J., et al. 2008, Aerodynamic Analysis of a Generic Fighter with a Chine Fuselage/Delta Wing Configuration using Detached Eddy Simulation, AIAA 2008-6228, 26th AIAA Applied Aerodynamics Conference
 Gortz, S., 2005, Realistic Simulations of Delta Wing Aerodynamics using Novel CFD methods, Ph. D Thesis, KTH Aeronautical and Vehicle Engineering
 Solfelt, D. A., 2007, CFD Analysis of a T-38 wing fence, Thesis, Department of the Air Force University