Preliminary design of a supersonic Short Takeoff and Vertical Landing (STOVL) fighter aircraft

w 5.2 CANNON RECOIL INDUCED YAWING MOMENT ESTIMATION . . 393 m r r w ¥ LIST OF SYMBOLS

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Definition

A Aspect ratio A g/ft/sec Root mean squared gust-induced g-level i.

A1 Aluminum ft.

b Wing span w.

Buttock line B.L.

ft.

Chord C C L Airplane lift coefficient Airplane drag coefficient

C_D

C_M Airplane pitching moment

w Parasite drag C_Do i.

C_Du Variation of drag coefficient with speed

I/rad

Variation of drag coefficient with angle C D_ot of attack w.

Variation of lift coefficient with speed C_L_u

I/rad

C_L_a Airplane lift curve slope Oi Variation of lift coefficient with rate C_L_a dot of change of angle of attack Variation of lift coefficient with pitch

C_Lq

rate C_M_u Variation of pitching moment coefficient with speed

l/tad

Variation of pitching moment coefficient C_M_cx with angle of attack vi C_M_o_ clot Variation of pitching moment coefficient with rate of change of angle of attack Variation of pitching moment coefficient C_M_q with pitch rate Imul Variation of side force coefficient with C._y__ sideslip angle Variation of side force ccefficicnt with C.j_r yaw ram Variation of side force coefficient with C_y_p roll rate Variation of side force coefficient with I/rad C_y_SR rudder angle 1/rad C_n_B Variation of yawing moment coefficient with sideslip angle C_n_r Variation of yawing moment coefficient with yaw rate Variationof yawing moment coefficient C_n_p with roll rate I/tad C_n_SR Variation of yawing moment coefficient with rudder angle I/rad C=.n_SA Variationof yawing moment coefficient with aileron angle I/tad _1_ Variationof rolling moment coefficient with sideslip angle Variationof rolling moment coefficient C_l_r with yaw ram Variationof rollingmoment coefficient C_Up with roll ram

Variationof rolling moment coefficient I/rad

C_LSR

with rudder angle l/tad C_I_SA Variation of roRing moment coefficient with rudder angle vii Cu Copper D Drag e Oswald's efficiency factor F.S.

Fuselage station o.

Horizontal tail H(subscripO Moment of inertia about x-axis slug-_ I xx slug-f_ Moment of inertia about y-axis I_.yy Moment of inertia about z-axis slug-_

Lzz

Product of inertia slug-f t2 I_xz Stabilator incidence dcg

LH

Correction factor K ft-lbs L Rolling moment w ft-lbs Perturbed rolling moment I/see a Dimensional variation of rolling moment L l3 about X s with sideslip angle I/sec Dimensional variation of rolling moment L_p about X_s withrollrate llscc Dimensionalvariation of rolling moment L_r about X_s with yaw rate llsc& Dimensional variation of rolling moment L_SA about X_s with aileron angle

l/scc_

Dimensionalvariation of rolling moment L_SR about X_s with rudder angle Mach number M M Pitching moment m Pernn'l_ pitching moment Mark Mk moment 1/ft=se.x_ Dimensional variation of pitching S_u withspeed Dimensional moment l/f t-see M_T_u variation of pitching withspeed moment 1/see a Dimensional variation of pitching with angle of attack 1/see a Dimensional variation of pitchingmoment M_T_o_ with angle of attack I/scc M a._dot Dimensional variationof pitchingmoment with rateof change of angle of attack 1/scc Dimensional variationof pitching moment M_q with pitch rate 1/s¢c _ M_SE Dimensional variationof pitchingmoment with elevator angle So Molybdenum ft-lbs N Yawing moment ft-lbs n Perturbed yawing moment Nautical miles nm moment about l/see a N_B Dimensional variation of yawing Z_s with sideslip angle moment about l/see a Dimensional variation of yawing N_T_B Z_s with sideslip angle moment about I/scc Dimensional variation of yawing N_p Z_s with roll rat.

moment about 1/see Dimensional variation of yawing N..r Z_s with yaw rate moment about 1/se_ N_SA Dimensional variationof yawing Z_s with aileronangle moment about 1/sec= N_SR Dimensional variationof yawing w Z_s with rudder deflection ra_sf_ P Roll rate ix Perturbed roll rate rad/$PA_ P Pitch rate rad/scc

Q

rad/scc Pcrmrbc(i pitch rate q Dynamic pressure lb/sq.R.

q_bar

R Yaw rate rad/scc

rad/s_ 1" Perturbed yaw rate S Planform area sq.R.

Ti Titanium lbs Tr_ Thrust required f_/scc U Forward velocity

ft/scc

U Pcm)xbcd forward velocity

V ft/scc

Side vclocity

ft/scc

V Perturbed side velocity Vanadium V Verticaltail V(subscript)

W ft/scc

Downward velocity

_scc

W Perturbed downward velocity W.L. Water line 1/scc X_u Dimensional variationof X_s forcc with speed

Dimensional variationof X_s force with 1/scc

X_T_u speed

fqsee 2

X_(_ Dimensional variationof X_s force with angle of attack

_sec a

X_SE Dimensional variation of X_s force with elevatorangle X ft/sec a Dimensional variation of Y_s force with Y_B sideslip angle ft/s¢_ Dimensional variation of Y_s force with Y..p roll rate ft/sec Dimensional variation of Y_s force with Y_r yaw rate r _ ft/sec _ Y_SA Dimensionalvariation of Y_s forcewith aileron angle v ft/sec a Dimensional variation of Y_s forcewith Y_SR rudder angle I/sec Dimensional variation of Z_s force with Z_u speed ft/scc 2 Dimensionalvariation of Z_s forcewith Z_ot angle of attack ft/scc Dimensional variation of Z_s force with Z_a dot rate of change of angle of attack ft/sec Dimensional variation of Z_s force with Z_q pitch rate Dimensional variation of Z_s force with f_se_3 2 Z__E elevator angle GREEK SYMBOLS O_ Angle of attack deg B _g Sideslip angle Deflection angle dcg w t.g) rad/scc NaturalFrequency v Damping Ratio xi AE Aerospace Engineering US dollars AEP Average Estimated Price AFCS Automatic Flight Control System w AGARD Advisory Group for Advanced Research and Development for NATO Air-to-Ground Missile AGM AIM Air Intercept Missile Annular Invertor Valve AIV AMRAAM Advanced Medium Range Air-to-Air Missile APU Auxiliary Power Unit ASRAAM Advanced Short Range Air-to-Ah" Missile AWACS Airborne Warning And Control System BAe British Aerospace BAI Battlefield Air Interdiction Best Cruise Mach and Altitude BCM/A BIT Built In Test BVR Beyond Visual Range CA Counter Air CAD Computer Aided Design ECM Electronic Counter Measures ECS Envu" onmental Control System EFA European Fighter Aircraft FLIR Forward Looking Infrared FOD Foreign Object Damage L xii HARM High Spe_ Anti-Radiation Missile -- HFVT Hybrid Fan Vectored Thrust -- HGR Hot Gas Reingestion -- Hands On Throttle And Stick -- HOTAS HUD Head Up Display -- IFF Identification Friend or Foc --

Ima g Infrared --

ILS

Instrument Landing System =-

IR Infrared -- ITR Inlet Temperature Rise =- JFS Jet Fuel Starter -- KEAS Knots Equivalent Airspe_ knots KU Kansas University US dollars LCC Life Cycle Cost LIFT Lift + Lift/Cndse Engine Cycle -- LTE Launch-To-Eject Cycle -- Mixed How Vectored Thrust -- MFVT MRM Medium Range Missile -- NACA National Advisory Committee on Aeronautics -- NASA National Aeronautics and Space Administration -- NATO North AtlanticTreaty Organization -- F_ NBC Nuclear, Biological, Chemical gear -- OBOGS On-Board Oxygen G-encrating System -- RCS Reaction Control System -- Reaction Control Valve -- RCV *l° gill RDTE Research, Development, Test and Evaluation RF Radio Frequency (Radar) RWR Radar Warning Receiver S&A Safe and Arm SAS Stability Augmentation System SLS Sea-Level Standard SRM Short Range Missile STOL Short Take Off and Landing STOVL Short Take Off and Vertical Landing TACAN Tactical Air Navigation Total Quality Management

TQM

UHF Ultra High Frequency v USD United States Dollars USRA Universities Space Research Association VHF Very High Frequency VTOL VerticalTake Off and Landing V/STOL Vertical/Short Take Off and Landing xiv 1. INTRODUCTION The survivability of long, hard surface runways at Air Force Main Operating Bases is fundamental to the current operations of the Air Force Tactical Air Command. Without the use of these runways, the effectiveness of the Tactical Air Command is severely degraded. One possible solution to this runway denial situation is to include a Short Takeoff and Vertical Landing (STOVL) capability in a supersonic fighter/attack vehicle.

Design teams at the University of Kansas, through the sponsorship of the NASA/USRA program, have completed a conceptual design study of three supersonic STOVL aircraft and based on this study, sekcted one aircraft for detailed design work.

The cooperation between the NASA/USRA Advanced Design Program and the design efforts at the University of Kansas are discussed in Section 1.1. Section 1.2 presents the study plan and objectives of the design study.

1.1BACKGROUND w The NASA/USRA Advanced Design Program is, from Reference 1.1, "a unique national program that brings together NASA engineers with students and faculty from United States engineering schools by integrating current and future NASA space/aeronautics curriculum." The University of Kansas is one of approximately forty five universities selected for this program.

The USRA Advanced Design Program course is taught in addition to the existing design courses at the University of Kansas. Table 1.1 shows how the USRA and KU design courses are offered to the students. Each design course is worth 4 hours of engineering design. All students are required to take AE 521 to learn the basic methods of design. The student is free to choose among the remaining five design courses to fulfill the eight hours of design required for a degree in Aerospace.

Although the USRA design courses are offered as a graduate level course, most students are undergraduates that wish to have more than the required amount of engineering design hours. Section 1.2 discusses in more detail the USRA design courses for the 1989-90 academic year.

1.2 STUDY PLAN = The supersonic STOVL started in the fall semester (Phase I) with a brief historical survey of powered lift vehicles followed by a technology assessment of the latest supersonic STOVL engine cycles under consideration by industry and government in the US and UK. A survey of operational fighter/attack aircraft and the modern battlefield scenario were completed to develop, respectively, the performance requirements and mission profiles for the study. Three aircraft were selected for initial investigations. The following engine cycles were used: a hybrid fan vectored thrust cycle, a lift+lift/cruise cycle, and a mixed flow vectored thrust cycle. Chapter 2 shows the results of the Phase I aircraft study. Chapter 3 presents the Phase I aircraft comparison and selection of the aircraft for the second semester (Phase _, in which the lift+lift/cruise aircraft was selected detailed design work.

Table 1.1 USRA and KU Desig'n Courses Fall Semester S t_ing Semester AE 522: KU Aircraft Design AE 521: Aircraft Design Detailed Analysis * Preliminary Analysis * Individual Work Team/Individual Work National Competition AE 523: Engine Design * Detailed Analysis * Team Work * National Competition AE 524: Space Design * Detailed Analysis m * Team Work USRA AE 622: Aircraft Design AE 621: Aircraft Design * Preliminary Analysis * Detailed Analysis * Team Work * Team Work Chapter 3 also discusses the design changes of this aircraft. Chapter 4 gives the aircraft description. The weight and balance is presented in Chapter 5 and the propulsion system integration is shown in Chapter 6. A takeoff, hover, and transition analysis is given in Chapter 7. The performance and mission capability of the aircraft is presented in Chapter 8. Chapter 9 presentsthe stability and control of the aircraft. Chapter I0 presentsthe material selection and structural layout of the aircraft and discusses asccssibility and maintainability considerations, including the engine removal. The aircraft systems layout is given in Chapter II. The weapons integration is shown in Chapter 12. The life cycle cost is shown in Chapter 13. Chapter 14 gives conclusions and recommendations for the study.

REFERENCES FOR CHAPTER 1 1.1 NASA/USRA University Advanced Design Program, Pro m'am Handbook for Facul .ty Teaching Assistants & Students, 1989-90 Academic Year.

2, PHASE I AIRCRAFT DESCRIPTIONS The purpose of this chapter is to present the results of the Phase I aircraft study. For each aircraft the following is given: a description of the configuration, a three view with geometric data, an inboard prof'fle, an area rule of the configuration, and a weight summary. Complete documentation of the Phase I aircraft are in References 2.1-2.3. The mission profile and specifications for the Phase I study are given in Section 2.1. Section 2.2 presents the Lift + Lift/Cruise configuration (the Monarch), Section 2.2 presents the hybrid fan vectored thrust configuration (the Viper), and Section 2.3 presents the mixed flow vectored thrust configuration (the Nemesis).

2,1 PHASE I AIRCRAFT MISSION PROFILE AND SPECIFICATIONS Reference 2.4 states the responsibilities of a fighter/attack aircraft in the European theater as a balance between counter air and close air support/battlefield air interdiction.

The design team chose to study a fighter aircraft having a primary mission as counter air and a secondary mission as battlefield air interdiction. The intent is to have the counter air mission size the aircraft with the battlefield air interdiction mission as a fallout.

The selected profile and specifications of the counter air mission and the battlefield air interdiction mission were developed using References 2.5 and 2.6. Reference 2.5 contributed the following information: the battlefield scenario for a STOVL fighter, the threats to a STOVL fighter (land based anti-aircraft and aircraft threats), research of similar aircraft mission profiles, and the stores and ammunition selection. Reference 2.6 contributed a mission capability trade study. This study investigated the sensitivity of aircraft weight to mission range and Mach number. Figures 2.1 and 2.2, respectively, show the counter air (CA) mission and battlefield air interdiction (B/d) mission profiles. Three BAI missions were selected for the study. The mission profile is the same for each, but the ordnance carried varies. The specifications for the missions are given in Table 2.1.

The point performance data was selected with the suggestion of Reference 2.7 that reasonable performance for a STOVL fighter can be selected by slightly bettering the performance of the Northrop F-20 (Reference 2.8). The selected performance is also shown in Table 2.1.

2.2 LIFt + LIb"r/CRUISE AIRCRAFT DESCRIFFION (MONARCH) The overall configuration of the aircraft consists of a conventional wing and fuselage with a canard and snake. The crew consists of one pilot. Payload requirements are given in the mission specification. The Monarch aircraft employs an unconventional internal mounting system for the counter airmission weapons. The engine cycle consists of one dedicated lift engine in the forebody of the aircraft and a lift/cruise engine in the aft end of the fuselage. The landing gear is of the tricycle type. A three view of the Monarch aircraft, including its geometric parameters, is shown in Figure 2.3. The inboard profile of this aircraft is shown in Figure 2.4.

Major design considerations for this aircraft include: * volume requirements for internal weapons,

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w I ORIGINAL PAGE IS OF PO0_ QUALITY Table 2.1 Phase I Aircraft Mission S__..cification CREW: One Pilot, (225 Ibs) w ARMAMENT: One internal M61AI Vulcan cannon, and 400 roundsof 20ram ammo Counter Air PAYLOAD: Two ASRAAM's (stored internally), and Two AMRAAM's (stored internally) Battlefield Air Interdiction Six Mk 82 Bombs (externally store_), or Four AGM-65 Mavericks (externally stored), or Six AGM-88A HARMs (externally stored) PERFORMANCE:

Performance Characteristic V_ue

Time to Climb 40k in 2 minutes lg Specific Excess Energy 500 ft/sec (2A) 30k 0.9M (2B) 10K 0.9M 1,000 ft/sec Sustained Turn Rate (3A) 0.SM/15k ft

15 deg/sec

(3B) 0.9M/30k ft

9 deg/sec

(3C) 1.2M/30k ft

8 dcg/sec

(3D) 0.9M/15k ft 6.5 g (3E) 1.6M/30k ft 4.5 g Acceleration (4A) 30k ft 0.9M to 1.6M 70 scc (4B) 0.5M to 1.4M 80 scc (4C) 10k ft 0.3M to 0.9M 22 scc Landing Distance Without Chute

2,2OO ft

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ORIGINAL PAGE IS OF POOR QUALITY * engine sizing and balance considerations for hover flight, * and supersonic flight requirements.

The Monarch aircraft has a cantilever wing configuration to provide primary lift for the aircraft in cruise flight. Full span leading and trailing edge surfaces provide high lift and roll control. The wing is mounted mid-fuselage with a 37.8 degree leading edge sweep. The addition of a straketo the aircraft provides delayed wing stall at high angles of attack, additional fuel volume and structural support for weapon hard points.The airfoils for the wing arc 8% thick at the root and 6% at the tip. The empennage of the aircraft consists of a conventional vertical tailand forward mounted canard. The vertical tailis a singlefm and houses a rudder to provide directional control. This rudder consistsof two individual pieces with separate actuators. This was done to provide redundancy against battledamage.

Primary design considerations for the fuselage layout include the requirements for internal weapons and shaping to reduce wave drag. Unconventional sizing was required to create internal volume for the counter airmission weapons and the dedicated lift engine.

This lead to the lower fuselagebeing flatfor most of the aircraft length."Coke-bottling" was incorporated at the wing fuselage interface in an effort to improve the area ruling.

The results of area ruling (Mach= 1) for the Monarch aircraft arc presented in Figure 2.5. Area ruling is a method used for shaping a fuselage to minimize wave drag in transonic and supersonic flight. From Figure 2.5 it can be seen that the Monarch aircraft slightly exceeds the ideal (Mach = I) area rule model. Removing the fuselagecoke bottling in the vicinity of the wing may provide a more favorable area distribution.

However, such a design change would add wetted area. This may resultin a net drag increasewhich would negate the area rule improvement.

The engine cycle specified for the Monarch aircraft consisted of a Lift + Lift/Cruise system: a dedicated lift engine for hover and transition and a lift/cruise engine used for hover, transition, and cruise. Design considerations for sizing the engines included hover, supersonic flight, and transition from hover to horizontal flight.

The landing gear chosen for this aircraft is a retractable tricycle type. It consists of a nose gear and two main struts aft. The main gear rea'acts aft into a fuselage fairing along the lift/cruise engine. The nose gear retracts forward to a position underneath the cockpit.

The cockpit of the Monarch is sized for one pilot. An ejection seat, heads-up display and center control stick make the cockpit conventional in design for a small or medium sized fighter aircraft. The view from the cockpit was an important design consideration in the Monarch. Lack of visibility is detrimentalto aerial combat effectiveness where the first sightingis very important. The pilotof the Monarch fighter will have a view of 14.5 degrees over the nose of the aircraft and 5 degrees over the tail. View over the side of the cockpit is 52 degrees.

Table 2.2 gives the weight summary for the Monarch aircraft.

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MISSION

_J:t3 CA STXUCTLTI_ (7349) Fuselage 4043 Wing 1579 Tails - Ve.:.ical 333 - Canard 298 Landing Gear - Main 931 - Nose 165 r PROPULSZ3N (6235) Cruise Engine 4009 Lift Engine 647 Air induction 876 Fuel _!adder 474 w Fuel D_pinq 26 Engine Controls 43 S=ar:ing Sys:em 139 water injecuicn 21 F_X_D E_U_PYv--NT (4484) Flight Con%r:l 999 AvLcnics 1164 Electrical System 548 Air C3ndi:ioni_g 254 Oxygen System 17 m ._U 257 Furnishings 276 Gun and Provisions 630 Auxilzr? Gear, Pain_ 341 (690) Venur%l Noz:ie 300 P,C3 E_&i=men_ 390 1B758 !87S+ !STY+ !_7_a -O'_.L. . " __..-:,_--" WE'GET_ 225 22S 22E 22_ Crew 10308 10308 10308 10208 Tc,zz! Fuel Ar.-.,ament (1290) (3820) (3820) (4300) .%_.KA_M S Y_:__ .M Mk-a2' s Maverick' s 220 220 220 220 ._--.2o - 200 ,nds 33111 33111 33591 30581 T._OFF WEIGHT ORiG!NAL PAGE IS OF POOR QUALITY u 2.3 HYBRID FAN VECTORED THRUST AIRCRAFT DESCRIPTION (VIPER_ A three view of the Viper aircraft, including geometric parameters, is given in Figure 2.6 and the internal layout is shown in Figure 2.7. There are five major configuration related aspects that drive the Viper design. These are: *) the forward swept wing, *) the empennage and tail configuration, *) the armament location, *) the fuel volume, *) and the powerplant and engine/airframe integration.

The overallaspectsof the Viper design are discussed below.

The Viper is equipped with triple redundant, fly-by-wire flightcontrol system. The vectoringof the exhaust nozzles, forward and aft,are also computer controlledfor stability and to maximize performance in transition and hover.

Forward swept wings in supersonic fighter configurations offer some advantages when compared to conventional planforms. An important consideration is the improved pilot visibility over the sides of the aircraft. This aspect is particularly important during vertical operations as well as during combat. A forward swept wing may also produce a smoother Sears-Haack area distribution, giving betterwave drag characteristics in the supersonic regime. This is important for a STOVL design which should not compromise its capabilities while operatingin a conventional mode during supersonic cruise.

A forward swept wing configuration allows for a great deal of flexibility in terms of structural synergism. One advantage is that the wing main spar frame is used to attach the engine to the restof the airframe. Another advantage is thatthe front spar and the kick spar are attachedto improve the su'uctural integrity of the wing.

Another characteristic of forward swept wings that is am'active in fighter applications is that the wing root will stall before the tip, allowing for continued aileron controlat high angles of attack. Furthermore, forward sweep allows for the wing center of gravityto be very close to the airplanecenter of gravity, decreasing the need for longitudinal trim as fuel and payload are expended.

A drawback that needs to be considered in the forward swept wing concept is that it is prone to body freedom flutter. However, this can be solved through aeroelastic tailoring, such as is done in the X-29.

The use of twin booms, like the forward swept wing, also allows for a great deal of synergism. They provide wing bending moment relief and volume for fuel and weapon storage. According to Reference 2.9, twin booms can also tailor the configuration for low wave drag, while a certain degree of combat survivability and redundancy is added.

Some problems, however, are associatedwith twin boom designs. They are: * In long boom configurations, critical loads on the taillead to large boom cross sections.

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ORIC-INAL PAGE IS OF POOR qUALITY ]5 * Vibration and fatigue due to excessive noise of engine exhaust flow.

* Scrubbing drag from the engine exhaust impingement on the boom structure may be a problem.

The booms do, however, produce a shielding effect on the exhaust, reducing the infrared signature of the aircraft.

An aftswept inverted vertical tailis used for the Viper. This design aids in the stealth characteristics of the aircraft as well as act as a structural tie between the booms.

The control surface of a V-tail must perform both of the jobs of a conventional elevator and rudder. Since a forward swept wing configuration lends itself to inherent longitudinal instability, a fly-by-wire system is needed.

For both counter air and battlefidd air interdiction missions, a M61A1 20 mm cannon is used with 400 rounds of ammunition. The gun is located under the fuselage on the port side. The counter air weapons are carried internally. The AMRAAM's are located in fairings at the wing root/fuselage intersection. The ASRAAM's are stored inside the booms. The battlefield air interdiction weapons am carried externally. There am six hardpoint locations to provide for this weapon capability: * two beneath the fuselage, between the nose and main landing gear doors, * two beneath the boom, where the boom intersects the wing, * and two beneath the wing, outboard of the boom intersection.

The Viper uses a hybrid fan vectored thrust(IIFVT) engine. The hybrid fan vectored thrust engine comprises a mixed augmented turbofan driving a remote fi'ont fan through a shaft. The I--IFVT has a dry thrust split of 0.6. The front fan is connected to the restof the engine by an interduct, at the forward end of which is a diverter valve.

There are two operating modes: I.

Parallel-- The front fan flow is divertedto a plenum and fed to two unaugmented, fully vectoring front nozzles. The core air is fed by a ventral auxiliary inletbehind the cockpit. The rear nozzle is vectorableto 110 degrees.

.

Series -- The auxiliary inlet and front nozzles are shut off with an annular inverter valve (AIV) that performs the miracle of flow shifting. The front fan air passes through the valve to the restof the engine. This provides for maximum engine boost.

The parallel mode is used in short take off,vertical landing and subsonic cruise. In short take off,the two front nozzles and the main rear nozzle are both vectored down and aft to createa lifting force and a forward velocity. All nozzles are vectored down during vertical landing. Using the parallel mode in subsonic cruise with the front nozzles vectored fullyaft will allow for a higher bypass ratio. This may, consequently, improve the specific fuel consumption in subsonic cruise.

The seriesmode is used for high performance and supersonic flight. In this mode, the front nozzles will be faired in by a retractable ramp to minimize drag.

= m A chin inlet is implemented in the Viper design. This is done for several reasons: * moving the inlet as far as possiblefrom the exhaust nozzles will reduce the hot gas re-ingestion (I-IGR) and foreign objectdamage (FOD), * thisposition allows for good pressure recovery, * and pilotvision is not affected.

The forward vectorable nozzles are located on the sides of the aircraft, just forward of the wing. This will allow for some lateral control by differential vectoringof these two front nozzles. However, a reaction control system (RCS) will still be requiredfor complete control.

Although the problem has not yet been thoroughly investigated, it will be assumed at thispoint thatengine removal will be accomplished by removing itout of the back of the aircraft. The structural arrangement of the Viper has not yet been determined, but engine removal will be a major concern. The very large front fan dimension may not allow for removal through the tail. As mentioned previously, the engine will be mounted to the wing main spar frame for structural synergism.

The largest contributor to drag in a supersonic flight regime is wave drag, often influencing the overall layout of an aircraft by dictating its cross sectional area distribution.

The area distribution for the Viper is shown in Figure 2.8 along with the ideal Sears-Haack Type I and II curves. The Viper matches the Sears-Haack Type II curve well along the forward fuselage,except for the canopy. Good visibility dictates this irregularity. A large increasein cross-sectional area occurs where the wing and wing glove begin. Because the u glove is relatively large,itvirtually counteracts the favorablegradual area build up of the forward swept wing. This is an aspect that should receive furtherconsideration in a future report. Coke-bottlingthe fuselage at thislocationmay decrease the effect.

The maximum cross sectional area is attained at roughly the midpoint of the aircraft. According to Reference 2.10, this maximum should occur between 55-60% of the aircraft length. The cross-sectional area decreases rapidly along the aft portion of the fuselage which is undesirablefrom a wave drag point of view. The irregularity thatoccurs as a resultof the empennage could be reduced through local coke-bottling of the booms.

Table 2.3 gives the weight summary for the Viper aircraft.

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MISSION

CX lUUll aa/t2 STRUCTURE (9055) Fuselage 5253 w Wing 2065 Tails - Horizontal 324 - Ve_ica! 996 Land/ng Gea= - Main 996 - Nose 176 PKOPULS_0N (6802) Cruise Engine - includes nozzles Air Induc:ion 475 Fuel Bladder 490 Fuel Dumping Engine Controls 22 189 .

S_artinq $ys:em Wazer injec:ion F_X_D E_U_P.__NT (4591) F!ighu Conuro!

Avionics 1!64 Eleczrical Sys:em Air Conditioning Oxygen System Furnishings Gun and _rcvisi0ns 630 Auxilary Gear, Pain_ ST0%_ ECUI_M_--NT (423) RCS Equipment TOT_ EM_-TY W_iGHT 20871 20871 20871 20871 Crew 225 225 225 225 To_a! Fuel 10754 10754 10754 10754 A.-mamen_ (!290} (3820) (3820) (4300) ASKAAM.S 400 _S 670 M.A2_M M.k-82' s Maverick's 4080 Ammo - 200 z-ads 220 220 220 220 TAKEOF_ WEIGHT 35670 35670 3_140 36150 r w 2.4 MIXED FLOW VECTORED THRUST AIRCRAFT DESCRIPTION (NEMESIS) A three view of the Nemesis aircraft, including geometric parameters, is given in Figure 2.8 and the internal layout is shown in Figure 2.9. The major aspects of the Nemesis configuration are discussed below.

The pilot's eye position is located to provide adequate visibility over the nose and sides of the aircraft. Additionally, the upper fuselage is carefully developed to avoid pilot "blind spots" behind the aircraft.

The large ducts needed for hover with the MFVT concept dictated the middle and aft fuselage width. The cockpit and radar sized the forward fuselage. Fuel volume considerations and the need for a long internal weapons bay for the AMRAAM sized the fuselage length. Volume beneath the engine inlet and ducts was dedicated to the main landing gear and ASRAAM missile storage.

Simple normal shock inlets were selected and sized to the Mach 1.6 supersonic dash requirement. A bifurcated inlet was selected so that the wide aft fuselage could be easily blended into the outside edges of the inlets. A chin inlet, e.g. F-16 Falcon, was not selected so as to avoid hot gas re-ingestion and FOD problems. The flat underside fuselage that developed from this integration should be beneficial in enhancing the fountain effects d_ring hover.

A conventional aft swept wing was selected for the Nemesis. This was done so that a simply constructed wing with adequate performance could be developed. Strakes have been incorporated to improve aircraft lift and to maintain adequate airflow to the bifurcated inlet at high angles of attack.

A tail aft configuration was selected for the Nemesis. This was done to keep the aerodynamic center near its originally estimated location, above the hover thrust location.

Additionally, the MFVT propulsion system had already created a wide aft fuselage with adequate structural allowances for all moving stabilators. Twin vertical tails were selected to provide adequate directional stability throughout the flight envelope.

The area ruling plot for the Nemesis appears in Figure 2.11. The constant cross- sectional areas of the fuselage ahead of the wing and in the vicinity of the propulsion system kept the area distribution of the configuration from matching the ideal Sears-Haack shapes.

A weightsummary for the Nemesis isshown in Table 2.4.

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MISSION

CA _tJ:#l a,_#2 _,I#3 STRUCTURE (7982) Fuselage 4237 Wing 1741 Tails - Horizontal 302 - Ver_ica! 379 Landing Gear - Main 1150 - Nose 173 PRCPL%S:CN (5983) Cruise Engine 4728 Air induc:ion 536 Fuel Bladder 484 Fuel D_T, ping 26 Engine Conurols 26 S:ar:ing Sysuem !_2 We:at _njeczion 21 F:XZD E_UZ_._-'_-NT (4_0) F!igh: Con_ro! i033 Avionics "!64 55i Eieczrica! System 2_4 Air C3ndi:ioninq Oxygen System _2U 272 27_ Furnishings Gun and _rovisions 630 3_3 Auxilar7 Gear, Pain_ ST0_% ECUI23'_NT (!36_) Kemoue !if: ducts insu!a:icn "60 Block and Turn Nozzle Clamshell Nozzle "C'.AL .-,a-_,-v _GKT I_B90 i}-:}0 19890 19890 Crew ...'s=" 225 225 225 Total Fuel 13575 10_r_ 10575 I0575 .%rm_enz (1290) (2_2_) (3820) (4300) A_._3_S 2 "CO -¢ MI<-82' S Maveri:k's Ammo - 200 :rids 220 22_ 220 22_ T._<ZCF_ WE:GHT 31980 34510 34510 34990 ORIGIN._,L PAGE _S OF POOR QUALITY REFERENCES FOR CHAPTER 2 2.1 Cox, Brim, et. al., Preliminary. Design of a Supersonic STOVL Aircraft Using a Lift + Lift/Cruise Engine Cycle, University of Kansas, AE 621, December 1989.

2.2 Cox, Brian, et. al., Preliminary. Design of a Su_rsonic STOVL Aircraft Using a Hybrid Fan Vectored Thrust Engine Cycle, University of Kansas, AE 621, December 1989.

2.3 Cox, Brian, et. al., Preliminary_ Desima of a Supersonic STOVL Aircraft Using a Mixed Flow Vectored Thrust Engine Cycle, University of Kansas, AE 621, December 1989.

2.4 Park, P.H., Fouga, S.G., and Kitowski, J.V., "Synthesis of the Next USAF Fighter - An Overview of Possible Requirements For an Advanced Multirole Fighter", AIAA/AHA/ASEE Aircraft Design, Systems and Operations Conference, Seattle, WA, August, 1989.

m 2.5 Cox, Brian, et al, Battlefield Arena in 1995-2000. University of Kansas, AE 621, September, 1989.

2.6 Cox, Brian, et al, Preliminary Performance Assessment for a Su__rsonic STOV'L i_, University of Kansas, AE 621, September, 1989.

2.7 USAF Wright Research and Development Center Presentation, Presenter: Second Lieutenant Gerald A. Swift, University of Kansas, October, 1989.

2.8 Hunt, B., and Titiriga, A., _, AIAA-86-2612, Northrop Corporation, Hawthorne, CA, October, 1986.

2.9 Dollyhigh, S.M., et al, Development and Analysis of a STOL Supersonic Cruise Fighter Concept, NASA Technical Memorandum, 85777, March 1984.

2.10 Nelson, B.D., Design Scope for Student Supersonic Projects, AIAA/AHS/ASEE Aircraft Systems, Design and Technology Meeting, October, 1986.

m 3. PHASE I AIRCRAFT COMPARISONS AND SELECTION OF PHASE II AIRCRAFT The purpose of this chapter is to compare the three STOVL aircraft from Phase I of the study and based on this comparison select the aircraft for Phase II. The aircraft are compared using the following parameters: 3.1 Aircraft Weights and Cost, 3.2 Aircraft Performance and Mission Capability 3.3 Area Rule and Drag Characteristics 3.4 Aircraft Components Required for STOVL Capability The selection of the Phase II aircraft is discussed in Section 3.5.

It is important to note that the configurations presented here are not converged designs due to lack of time in the Phase I study. Nevertheless, it is felt that the comparisons made here are still valid for preliminary design purposes.

3.1 AIRCRAFT WEIGHTS AND COST A comparison of the Phase I aircraft weights for the counter air mission am given in Table 3.1.

Table 3.1 Phase I Aircraft Weight Comparison (Counter Air Mission) (all weights in pounds)

LIFT HFVT MFVT

WTo 30581 33140 31980 WE 18758 20871 19890 W,_,, 7349 9055 7982 = Wp,_ 6235 6802 5983 W_ 4484 4591 4560 WF 10308 10754 10575 The HFVT aircraft is the heaviest due to its propulsion system and the boom arrangement.

The LIFT aircraft is the lightest configuration which is consistent with data presented in Reference 3.1.

The aircraft cost is summarized in Table 3.2. These cost estimates, in 1995 dollars, are based on 1,000 aircraft operating 350 flight hours per year for 20 years.

w T_ble 3.2 Phase I Aircraft Cost Comparison (Billions of 1995 Dollars)

LIFT HFVT

59.42 59.72 59.26

Life Cycle Cost

3.45 3.45 2.84 Research, Test, Development and Evaluation 19.76 20.21 Acquisition Cost 19.46 m 35.91 35.62 Operating Cost 35.91 0.60 0.59 w Disposal Cost 0.60 Cost per Aircraft 23.20 23.10 (millions) 22.90 w Although LIFT aircraft is the lightest configuration, its added cost for the lift engine makes its cost as much as the MFVT and HFVT configurations.

3,2 AIRCRAFT PERFORMANCE AND MISSION CAPABILITY The performance requirements from the missions specifications were verified and are shown for the three aircraft in Table 3.3. The aircraft meet the required performance except for the time to climb and specific excess energy for the LIFT and HFVT configurations. The lack of adequate performance shown is due to optimistic estimation of the wave drag in the preliminary sizing of the aircraft. The MFVT aircraft, which requires dry thrust for vertical operations, met the requirements since its engine was oversized for hover.

The mission capability was measured by estimating the fuel required to meet the design missions (see Figures 2.1 and 2.2). Table 3.4 shows the mission fuel burn for the aircraft. All three configurations can meet the mission ranges with the MFVT aircraft using the least amount of fuel.

Table 3.3 Phase I Study Aircraft Point Pm'formn.eeS Requirement H M Value LIFT HFVT MFVT 2 min 2.51 0 0 to 40k, 2.18 2.00 30000 0.9 500 ft/sec 417 516 i0000 0.9 1000 ft/sec 933 980 15O0O 0.8 15 16.34 16.19 deg/sec 17.90 30000 0.9 9 9.22 deg/sec 9.85 10.60 30000 1.2 8 10.10 deg/sec 7.92 7.80 30000 1.6 4.5 8.67 6.88 g 6.50 15000 0.9 6.5 8.25 8.18 9.60 g 30000 0.9 to 1.6 in 70 sec 40.20 38.10 34.70 30000 0.5 to 1.4 in 80 sec 57.70 52.40 49.30 I0000 0.3 to 0.9 in 22 sec 17.80 16.80 16.10 Table 3.4 Phase I Aircraft Mission Fuel Burn

LIFT

MFVT

CA Mission 7509 Ibs 8062 lbs 6917 lbs BAI Mission 9695 Ibs 10299 lbs 7995 lbs 3.3 AREA RULE AND DRAG CHARACTERISTICS A The area rule plots for the three configurations were shown in Chapter 2.

comparison of the area rule and drag characteristics is given in Table 3.5.

Table 3.5 Phase I Area Rule and Drag Chamcteristic_

LIFT I-IFVT MFVT

Match with $ears-Haack Fair Unacceptable Unacceptable Maximum Area 4073 in = 4709 in = 6303 in = 0.016 Wave Drag Increment at M=l.6 0.012 0.016 Aircraft Skin Friction Coeff. 0.0036 0.0051 0.0030 (M-0.8, H--30000 It) The unconventional fuselage shaping for the propulsion systems of the HFVT and MFVT concepts caused unacceptable area rule plots and also large maximum cross sectional areas, both of which increase wave drag.

3.4 AIRCRAFT COMPONENTS REOUIRED FOR STOVL CAPABILITY The weight and volume for the components required for STOVL capability axe presented here. Table 3.6 shows the components required for each aircraft along with their weights and volumes.

Table 3.6 Weight and Volume Data for STOVL Comnonent._ Volume (ft^3) Weight (ibs) LIFT * Lift Engine 21 647 * Ventral Nozzle and * 300 Turning Vanes * RCS System 8 390 Total 29 1337 HFVT * Flow Switching Mechanism 83 1351 and Extended Power Shaft 2 * * Front Vectoring Nozzles * Rear Vectoring Nozzle * Penalty for Booms 117 1112 * RCS System 6 423 Total 208 2886 MFVT * Block and Turn Nozzle * 450 * Transfer Ducts 92 465 * Front Clamshell Nozzles 2 450 D_w Total 94 1365 The HFVT configuration suffers the most from the STOVL equipment for two reasons.

First, the engine components required for flow shifting are heavy and require a large volume. Second, the engine thrust split requires the engine to be at the center of the aircraft and thus some sort of boom configuration. The LIFT and MFVT configurations have similar weight penalties but the MFVT has a larger volume penalty due to the transfer ducts.

3.5 SELECTION OF PHASE II AIRCRAFT The lift+lif_cruise configuration was selected for the Phase II aircraft study. The reasons for this selection were: 1) The LIFT configuration exhibited the most promising area rule distribution.

2) The technology required for this configuration is the most consistent with the 1995 Technology Availability Date (TAD) assumed for the study.

3) The LIFt aircraft was the lightest configuration.

At the start of the Phase II study, the LIFT configuration was iterated to reflect comments made about the design from References 3.1-3.3. The following were the drivers for the iteration: It is good if: * the aircraft center of gravity moves aft * the CA and BAI mission cg's in hover are aligned * the rear thrust post is moved forward * the aircraft has three posts instead of two * the lift engine is small With these considerations, the following modifications were made to the design: 1) The aircraft has a horizontal tail, not a canard. The purpose of this iteration was: * to move the cg further aft * to reduce the complexity in the main inlet region * to have more favorable stability margins 2) The aircraft has thmc postsinstead of two. This was done to allow for.

* reducedsuckdown in ground effect * hoverroll control throughdifferential areachange 3) The wing was shifted forward I0 inchesto achievea smaller positive stability margin in supersonic flight.

4) The avionics were moved aft in the aircraft behindtheinternal weapons bay to move the hover cg rearward, thus decreasing the size of the lift engine.

5)

The BAI mission payloadswere changed to reflect more realistic missions according to Reference 3.1and 3.3. The mission payloads are now configured to allowcarrying r radarguided weapons (Mavericksand HARM's) along with unguided weapons (Mk 82), thus having the aircraft capable to deliver munitions even if the target shuts off its p radar.

The BAI missions (two of them) were changed to: * BAI Mission #I - Four Mk-82's and two HARM's * BAI Mission #2 - Four AGM-65 and two Mk-82's 6) The short range missiles were placed on the wing tip for two reasons: * the target field of view of the missile is greatly enhanced * the missile must have "lock-on" before it is hunched, and external carriage allows more operational fi_lom.

7)

The design missions were scaled down to get a more realistic fuel fraction according to Reference 3.1. The counter air mission was scaled down to a 100 nm subsonic cruise and a 50 nm supersonic cruise. The bardefield air interdiction mission was scaled down to a 200 nm subsonic high level cruise and an 80 nm low level dash.

8) Actual data of the General Dynamics F-16 and Grumman F-14 wave drag increments were used to estimate the wave drag of the configuration.

9)

The weights of the following components were adjusted based on previous industry and government aircraft studies and actual aircraft: * cruise engine * installed avionics * reaction control system * rear and ventral nozzles * internal weapons launching mechanisms 10) The landing gear was re-sized for soft ground capability.

11) The wing thickness ratio is 4.5 percent for more favorable area rule characteristics.

The result of these design modifications is described in Chapter 4.

r w w REFERENCES FOR CHAPTER 3 3.1 USAF Wright Research and D_velopn_nt Center l:_sentation, Prcsent_r: Ray Fredette, University of Kansas, January, 1990.

3.2 KU STOVL Presentation and Design Discussion, Presenters: Brian Cox and Paul Borchcrs, Wright Research and Development C_.cntcr, January, 1989.

3.3 NASA Ames Research Center Presentation, Presenter: Andrew Hahn, Powered Lift Technology Branch, Universityof Kansas, February, 1990.

4. CONFIGURATION DESCRIPTION The purpose of this chapter is to give the configuration description of the Monarch lift+lift/cruise supersonic STOVL aircraft. A three view of the aircraft with a table of geometric parameters is shown in Figure 4.1. The internal layout is shown in Figure 4.2.

The requirements that had a major impact on the Monarch design are: * the short takeoff and vertical landing capability, * the supersonic cruise and combat conditions, * and the internal volume for medium range missiles.

The Monarch configuration de,couples the short takeoff and vertical landing capability from the supersonic requirements by employing a lift+lift/cruise engine cycle.

The lift engine, sized for the hover flight condition, allows the mission performance requirements to size the lift/cruise engine, thus making the propulsion system integration of the Monarch a more conventional integration than other STOVL concepts. The Monarch has a pitch and yaw vectoring nozzle system to allow for enhanced maneuvering a post stall conditions and, in the yaw axis, to augment the directional control.

w The supersonic cruise and combat conditions required the Monarch to have a smooth area rule distribution that matched the ideal Sears-Haack shape. Figure 4.3 shows thatthe Monarch met thisrequirement.The internal volume required for the medium range missileswas offsetby the wing thickness selection. The Monarch uses a 4.5 percent thickness to chord ratio for its aft swept wing. The strakeon the wing was included to provide for delayed wing stall at high angle of attack and for vortex lift in maneuvering.

The empennage of the Monarch consists of a singlevertical tail and allmoving horizontal stabilators. The size of the vertical tailwas reduced and the rudder removed by using the yaw vectoringnozzle. The size and placement of the stabilators were selectedwith the desirefor the Monarch to have _al trim drag throughout the flight envelope.

The high inlet placement was the result of two requirements. First, a low inlet placement would have required the inlet to shape itself around the internal weapon bay which was not desirable since this would have distorted the flow. Second, a higher inlet placement leads to less severe hot gas reingestion and foreign object damage problems.

The Monarch carries two medium range missilesinternally and two short range = missileson the wing tips for the counter air mission. For the fallout battlefield air interdiction mission, the Monarch carriesa combination of guided and unguided munitions on wing pylons. Wing pylons were selectedfor two reasons. First, pylon mounted stores allowed more flexibility in maintaining a constanthover centerof gravity, which is _=_- important to the lift engine sizing. Second, stackingmunitions underneath the fuselage interfered with the internal weapons bay doors, eliminating the possibility of a combined counter air and battlefield air interdiction mission.

The landing gear tires for the Monarch are oversized for a fighter due to the fact that a STOVL type aircraft may often find itself in an austere battlefield scenario with soft w field landing and takeoff conditions.

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$ i W pm J w w (S'_-DN] _lVf_S) _ T_NORD_$ $SO_I3 S. WEIGHT AND BALANCE The purpose of this chapter is to present the Monarch weight and balance results.

The weight and balance method is first presented, foLlowed by the weight and balance data.

Figure 5.I shows the weight and balance flow chart used for the design. As shown, the three primary drivers for the weight and balance are having: * the hover cg and thrust center balanced, * the inflight cg travel acceptable, * and an acceptable static margin.

w Fimn_ 5.1 Weight and Balance Flow Char{ ORIGINAL PAGE IS OF POOR QUALITY A secondary driver is to assure that the weight data and placement of components are reasonable. The weight data were estimated using empirical weight equations of Reference 5.1 and actual weights from operational aircraft. The weight and balance calculations are shown in Appendix I. The finalweight statement for the Monarch is shown in Table 5.1.

The center of gravity excursion diagrams for the counter air and bardcfield interdiction missions are shown, respectively, in Figure 5.2 and Figures 5.3 and 5.4.

The int'Light center of gravity travel is within the acceptable range given in Reference 5.2.

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I _sEL-_G_ s'ri_o's _ _ I i, , 20OO0 4,40 4,_ 450 G0 46O I ) , .,! ) I ,1 0_.25 0.30 0.35 0.40 FRACTION OF W_G MGC F_.ne 5.2 Counter Air M_,_on GC Excur_on Dia_r-_m (See Table 5.1 for CA, BAI#1, and BAI#'2 Ordnance) -- Table 5.1 Weight S.mmary. of the Monarch Airum_ BAI#1 BAI#2 CA (9498) STRUCTURE Fuselage Wing Tails - Vertical - Canard Landing Gear - Main - Nose Launch Mechanims (Int. Weap) ASKAAM AMBAAM Ventral Clamshell Nozzles PKOPULSION (6139) Cruise Engine 48O Lift Engine Cruise Engine Tailpipe Ext Cruise Engine Nozzle Air Induction Fuel Bladder Fuel Dumping Engine Controls Starting System (5480) FIXED EQUIPMENT Flight Control Avionics Electrical System Air Conditioning Oxygen System APU Furnishings Gun and Provisions 630 Auxilary Gear, Paint RCS Ducting and Nozzles 21117 21117 21117 TOTAL EMPTY WEIGHT 225 225 225 Crew 8642 8642 Total Fuel Armament (1196) (4074) (3316) ASRAAMS AMBAAMS HARM 2240 1120 Mk-82's Maverick's 220 220 220 Ammo - 200 rnds 33642 TAKEOFF WEIGHT 31336 34400 38a

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, 1 , 1 t t L 0.2.5 0.30 0.35 0.40 FRACTION OF WING MGC Fimu-_ 5.3 BAI #I Mission CO Excu_.-sion D_m'm"n (S¢c Table 5.1 for CA, BAI#1, and BAI#2 Ordnance) 3600O _4000 32000 30000 20000 260O0 .< L_ == m < 22000 20000 FRACTION OF WilCO MGC Fimn'e 5.4 BAI #2 Mission CG Excursion Diam'a m (See Table 5.1 for CA, BAI#1, and BAIW2 Ordnance) 4O REFERENCES FOR CHAFI'ER 5 5.1 Roskam, Jan, Part V. Component Weight Estimation, Roskam Aviation and Engineering Corporation, Route 4, Box 274, Ottawa, KS, 66067, 1985.

5.2 Roskam, Jan, Preliminary_ Conf_m,_tion Dcsi_ and Intem-ation of the Prooulsion _.Ica_, Roskam Aviation and Engineering Corporation, Route 4, Box 274, Ottawa, KS, 66067, 1985.

6. PROPULSION SYSTEM INTEGRATION The purpose of thischapter is to describethe integration of theMonarch propulsion system. Section6.1 describes the cruiseengine and 6.2 describes the lift engine. Each section describes the engine as well as the inlets and nozzlesassociated with the engine.

Figure 6.1shows the completepropulsion system as itis integrated in the airframe.

6.1 CRUISE ENGINE The cruise enginethatisused isbased on an engine prodded by Reference6.1.

The engine isdesigned to operate in both the hoveringas well as the cruise/maneuver flight conditions. The following sub-sections willdescribe the engineas well as theinlets and nozzles.

6.1.1 Cruise En_ne Description and Perforrnanc_'.

The cruise engine was sized for both the hover and conventional wing-borne flight conditions. The total dry thrust required from the cruise engine during hover must be 1.30 times the weight of the aircraft in hover which is 24744 lbs. This factor is based on the following: 1) The total vertical thrust during hover must be sized to include the following factors: a) 1.0g is to provide a force to counter the weight of the aircraft.

b) 0.1g is to enable the aircraft to counter a tenth of a g sink rate.

c) 0.03g is for out-of-ground suckdown (assumed) d) 0.1g is for in ground effect suckdown (assumed) 2) The cruise engine must also be ableto support the RCS which is0.07g.

Based on these parameters, the rcquir_ thrust from the cruise engine is 19,800 lbs dry. A point performance determined that a takeoff thrust-to-weight ratio of 1.15 was required for a maneuver Right condition; therefore, the engine must produce 35,450 lbs of thrust augmented. This means that the maneuver condition is more critical and determines the size of the engine. The base enginewas resized using the following scaling laws from Reference6.2.

New Length = Base Length*0New Thrust/Old Thrust) °" New Radius = Base Radius*(New Thrust/OldThrust) °a New Inlet Airflow = Base Airflow*(New Thrust/Old Thrust) Table 6.1givesthe Monarch engineparametersand the engine dimensions are shown in Figure 6.2. The engine weight includes the engine, fuel and oil systems, gear box, necessary plumbing, and mounting hardware. The performance plots for the installed cruise engine arc shown in Figures 6.3 through 6.5 for three engine ratings: maximum augmented, maximum unaugmenteck and a partial throttle setting. Figure 6.3 shows the mass flow rate for the cruise engine at various altitudes and math numbers. Figures 6.4a t_

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Table 6.1: Cruise EnEine Parameters v Max Dry_ Thrust Max Aug. Thrust Condition SLS 90°F day SLS 90°F day Mass Airflow 319.64 Ibm/see 319.64 lbm/sec Nozzle Throat Area 3.431 ft a 5.268 Bypass Ratio 0.80 0.80 Nozzle Pressure Ratio 3.268 3.096 Net Thrust 24,673 Ibs 35,573 Ibs Diameter 44 in. 44 in.

Length 184 in. 184 in.

Weight 3557 lbs 3557 lbs SCALE 1/50 ALL DIMENSIONS INCHES

L 1

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Since the maximum operating speedof the aircraft does not dictate the use of a variable geometry inlet, a normal shock inlet isused,and as discussed in Chapter 11, a bifurcated inlet isbetter than a chin inlet for alleviating hot gas rcingestion. Therefore, a bifurcated normal shock inlet is used on the Monarch. According to themethods of Reference6.3, and usingthe engine data from Reference 6.1, the total capturearea of the bifurcated inlet is calculated to be 6.89fr _. This capnLreareasize isbased on the following assumptions.

areal The inlet sizing pointisthe supersonic operation at M = 1.6and 30,000 fl altitude.

The current engine mass flow rate for the given flight condition is 319.64 lbngsec.

The ratio of secondary air flow to engine air flow (Ms/Me) is assumed to be 0.2 (Reference 6.3).

The mass flow of the boundary layer bleed is 3 % of the inlet capture area (Reference 6.3).

A dimensioned front view of the irdct lipshowing the capture area and shape is shown in Figu._6.6 BL = -]B FS = 2SO ALL DIMENSIONS SCALE t/20 _N_ ARE INCHES LEFT INLET SHOWN.

RIGHT INLET IS THE w MIRROR IMAGE 32.4 I G.2 ----.-D J WL = I55 Fimn'e 6,6 Cruise Inlet Lip Shape w The inlet lips arc placed on each side of thefuselage as high as possible to avoid FOE) and HGR (sccChapter7). They arealso placedbehind the cock-pit forimproved pilot visibility. The exactlayoutof the inlet from the mouth to the compressor isdesigned to avoidconflicts with any systemswhile attempting to maintain the highest inlet efficiency. The inlet layoutisshown in Figure6.7.

A channel typeboundary layer splitter isused. According to Reference6.3, the width of the boundarylayer at the inlet can be assumed to be I% - 3% of thelengthof the fuselage ahead of the inlet. 2% isused fortheMonarch which results in a width of 5 inches.Thereforetheboundary layer splitter is placed5 inches from the fuselage.

It isnecessary to insurethatthe cruise engine has sufficient airflowatall times including low speed and hover flight conditions. The bifurcated inlet describexl must be designedforthesupersonic flight conditions which means thatitdoes not have sufficient capture area at thelow speed conditions. Therefore, auxiliary inlets willbe placeon top of the main inlets and willoperateonly duringthe low speed flight conditions. According _q to Reference6.3, theidealinlet duringhover is a bellmouth since there arc no ram effects.

Geometric consu'aints make thisimpossible, so itisassumed for preliminary designthat a capture areaof 1.15timesthe compressor diameterissufficient. The compressorarea is 10.56 ft a,which means thatthe total inlet captureareamust be 1.15timesgreater or 12.14 fla.As previously stated, the cruise inlet capturearea is 6.89 ft a. Therefore, the total auxiliary inlet capturearea is5.25 fla.

The location of the auxiliary inlets should be such thatthe airfrom them sufficiently mixes with the air from the main inlet openings beforereachingthe engine face. They shouldalsobe locatedsuch thatthe total airflow is accelerated to approximately Mach 0.5. The size, shape,and location of these inlets arcshown in Figure 6.6. A permanent screenwillbe placedover the auxiliary inlets to preventFOE). The lossesdue to thisscreen arc considered negligible due to the low speed. Based on Reference6.5, a setof horizontal louverswillbe over the inlets to sealthem during c_'uise flight and open duringlow speed flight. Louvers are viewed as being the easiest to w mechanically operateplusthey should actas flow _g veinswhen openccL Since they willonly be operated duringvery low speed, itis not believed thatthe most forward louverwillblockthe flow intothe aft louvers.An elcctromechanical actuator willbe used to operatethe louvers, and willbe placein theinlet boundary layersplitter. Figure6.7 m also shows thelocation of the auxiliary inlets as well as a schematicof the actuation.

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" / i" ,/ w 6.1.3 Cruise En_ne Nozzle,s The cruise engine has two types of nozzles.One nozzle isthemar nozzle providing thrust vectoring which is shown in Figure 6.11. The othertypeof nozzleis a pair of ventral nozzles prodding hover capabilities which am shown in Figure 6.8.

The designdriver forthe ventral nozzleswas thatthe nozzles must have variable areacapabilities along with thrust vectoring of 15 degrees aboutthe x-axisto allowfor translation. The ventral nozzles were sizedby getting the throat areafrom Reference 6.1 and converting itto an equivalent areaforeach ventral nozzle. Therefore, the ventral nozzleswillhave therequired throat area to kccp theflow "choked"as the rearnozzle blocksthe airflow.

The ventral nozzlesarc shown in Figure6.8. The clamshell nozzleis a low weight, a low complexity, and a variable area nozzle. The othernozzle considered was one with turning vanes. The primary problem with theturningvane nozzleis thattheflow must be vectoredto reducethethroat area. This isnot acceptable for theventral nozzlesbecause the nozzleswillbc used forroll control by differential thrust of the two ventral nozzles, which require variable areacapabilities withoutlossof thrust along the z-axis.The clamshell nozzleswillbe retracted for up and away flight. Fuselagedoors willbe used to reducedrag thatwould be caused by the exposed ventral nozzles.

The turning vanes, as shown in Figure6.8,helpto alleviate pressure losseswhen turning theflow 90 degrees.The sizing of the ductwas calculated assuming 5 percent lossin pressure in the duct. The ventral nozzleducts must be detachable from themain v engine so that expedient engine removal ispossible.

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The rearnozzle must providepitchand yaw vectoring capabilities as discussedin Chapter 4. The primarydriver for a pitchand yaw vectoring nozzleisto provid_ enhanced maneuveringcapabilities and allow for removal of therudder. The following rearnozzledesignswere looked at: 1) 2D convergent-divergent nozzle with 20 degree pitch and 15 degree yaw capabilities. The nozzle could also block and turntheflow,which is required for hover. The primarydrawback to thisnozzleisthe complexity.

The secondarydrawback isthatto produce a side forcethenozzlehad to bc spoiled (similar to thrust reversing, but not as extreme), causing largelosses in axial thrust.This nozzleis shown in Figure6.9.

2) An axisymmctric nozzlewith 20 degreepitchvectoring and block and turn _vwwwwwwww_ capabilities. This nozzle would have bccn used ff nozzle #3 did not produce theside-force requiredtoremove the rudder. The reasonthisnozzlewould have been used isbecauseof itslow weight relative to a 21) nozzle. This nozzleis shown in Figure6.10. Because nozzle#3 provides the required side-forces thisnozzlewas excluded.

3) 2D convergent-divergent nozzlewith 20 degree pitch and 25 degreeyaw capabilities. This nozzle can alsoblock and turn the flow,which isrequired for hover. The reasonthatthisnozzleisbetter than nozzle#1 isthat the yaw vectoring occurs after thenozzle. Therefore, the axial thrust lossis reduced. The drawback to thisnozzleisitssizeand weight arc larger than nozzle#2. Nozzles#3 and #I are similar in size and weight. This nozzle producesenough side-force to eliminate therudder as discussed in Chapter 10. This nozzleisshown in Figure6.11.

Nozzle #3 was chosen forthe Monarch because of the capability to remove the rudder as discussed in Chapter 9. The primarydrawback to nozzle#3 isthattheweight is 20% greater than theothernozzleoptions.Nozzle #3 isshown in Figure6.11.

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'" ; I LL I 6.2 LIFT ENGINE The lift engine used forthe Monarch isbased on theRolls-Roycedirect lift engine.

The engineparameters for thisengine are taken from Reference 6.5. Itis an unmixed turbofan designedto provide vertical thrust fora STOVL aircrafL The technologystandard assumed for this designis consistent for an initial operational capability of 2005.

According to Reference 6.5 the RoUs-Royce engine was designed for vertical mounting and includes a vectoring exhaust nozzle. The engine has a large amount of parts made with advanced composites, which enables the uninstalled thrust to weight ratio to reach 28. To achieve the lightest possible solution whilemaintaining acceptable jet exhaust conditions, a relatively high bypassratio isimplemented. A higherbypassratioresults in a higher engine volume. A smaller diameter engine with a higherspecific thrust could be used todecreasetherequired engine volume. However, thiswillleadto an increase in engineweight and/or more severeexhaustconditions.

6.2.1 En_ne Description and Performance The size required for thelift enginewas determined solely by thethrust requirements of hover. As mentioned in subsection 6.1.1, the total thrust required during hover is1.23timesgreater than thehover weight of the aircraft. The amount of thrust from thelift engine was determinedby balancingthethrust from both enginesabout the center of gravity of the aircraft. The thrust balance at hover for the Monarch is shown in Figure6.12. The original Rolls-Royce engine was resized usingthe same scaling laws that were used for the cruiseengine. The re'sized engine parameters are listed in Table 6.2.

Table 6.2: Entdne Paramet_l'8 Max Dry. Thrust Condition SLS 90°F day Mass Airflow 266.37 lbm/sec Nozzle ThroatArea 3.431 ft a Bypass Ratio 1.5 Nozzle Pressure Ratio 3.268 Maximum InstaLled Thrust 12,105 lbs Diameter 32.8 in.

Length 35.1 in.

Weight 480 lbs 6.2.2 En_ne Air Induction System The lift engine inlet is positioned at fuselage station 230. Due to the close proximity of the engine to the cockpit, a bifurcated inlet isused. Since this engineis only used duringhoverand transition, the total inlet capturearea isassumed to be 1.15times greater than the compressor area or 6.83 fla. Louvers will also be used to seal the inlet during wing-borne flight, and will operate similarly to the auxiliary inlets.

Figure 6.13 is a cross-sectional view of the lift engine and inlet including the louvers.

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II m m m Clb t',- m % M u II It II CROSS-SECTIONAL CUT AT FS 230 SCALE: 1/:30 200 WL CLOSED J OPEN LOUVERS / LOUVERS -I = w BL I I ! _ I 20 0 -20 Figure 6.13 Cross-section View of En_ne Inlet for the Monarch 6.2.3 En_ne Nozzle The lift engine nozzle was designed so that the thrust could be vectored 20 degrees forward and aft to allow for pitch control. The nozzle was designed so that during up and away flight the thrust vectoring vanes will close. Therefore, no fuselage doors are needed.

This nozzle design is shown in Figure 6.13. The lift engine nozzle vectoring vanes are powered by two electromechanical, jack-screw, actuators. The vectoring may allow the lift engine to enhance the pitch control of the aircraft during hover and transition.

A gimballing nozzle, similar to that used for a rocket, was considered for the lift engine. The problem with the gimballing nozzle is greater complexity than the vectoring nozzle and also the need for a fuselage door. Another nozzle considered for the lift engine was a clamshell nozzle, similarto the ones on the ventralnozzles. The primary drawback of the clamshell nozzle is that the clamshell nozzles occupy more volume than the thrust vectoring vanes. The reason the clamshell design is used for the ventralnozzles is that the ventralnozzles are required to be variable area nozzles.

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17.0_ miii; Lift EnL,'ine Nozzle for the Monarch Fi_s_ 6.14 REFERENCES FOR CHAPTER 6 6.1 Pratt and Whitney Advanced Engines Performance, Weights and Dimensions Model, Fighter Attack/Interceptor Application, Provided by Public Release from Wright Research and Development Ccn_r, January 1990.

6.2 1989 AIAA Sm_nt Coml_tidon Engine Pc..rformance Deck.

6.3 Raymer, Daniel P., Aircraft Desien: A Conceptual Approach, American Institute of Aeronautics and Astronautics, Washington D.C., 1989.

6.4 Swavely, C.E., "Propulsion System Overview, Turbine Engine Types/Characteristics/Technologies, United Technologies Pratt and Whitney, 6.5 Roils Royce, ASTOVL Direct Lift Engine for AIAA Stuclent Project; Performance and Installation Brochure, July 1989.

7. TAKEOFF. HOVER AND TRANSITION ANALy$I$ The purpose of this chapter is to analyze some of the unique features of STOVL aircraft during operation below the velocityfor wing-borne flight. The following topics are covered.

Section 7.1 TAKEOFF GROUND ROLL DETERMINATION Section 7.2 TRANSITION ANALYSIS Section 7.3 HOVER ANALYSIS Section 7.4 PILOT WORKLOAD ANALYSIS 7.1 TAKEOFF GROUND ROIJ_ DETERM]2q'ATION The following step-by-step procedure is used by the Monarch for short takeoffs.

Step I With the airplaneat the beginning of the runway, and the brakes on, the auxiliary inlets and the lift engine inlets are opened, the leading edge flap is deflected 20 ° and the trailing edge flap is deflected 40 ° . Then both engines are started.

Step 2 While keeping the brakes on, the Cruise engine is throttled up to maximum dry thrust and only the main nozzles is used. However, it is deflected 20 ° downward to balance out the moment created by the idlingLift engine.

Step 3 The brakes arc released,the aircraft begins to move, and the Lift engine is throttied up Step 4 When the airplane has acceleratedenough that the wing provides sufficient lift for the wing and engines to lift the airplane, 16,312 Ibs of thrust is diverted to the ventralnozzles of the Cruise engine. This thrustcombined with the 11,514 Ibs of thrust provided by the Lift engine and the wing will lift the airplane into the air.

Figure 7.1 shows the thrust vectors produced by the engines at critical stages of the takeoff as well as the equivalent thrast. The times and distances shown are for the Counter Air Mission. It should be noted that at all points during the takeoff the total thrust is balanced indepcndendy of the aerodynamic forces on the aircraft. Also, the aircraft has 2 ° of ground incidence but it does not rotate to takeoff. This was not desired since the ventral nozzles cannot be deflected aft; therefore, they would produce a component of drag.

As seen in Figure 7.1, following this procedure using precise thrust angles and magnitudes, the Counter Air Mission takeoff ground roll was determined to be 238 ft.

This distanceis determined using lift and pitching moments in ground effects and a friction coefficient of 0.2. The Lift engine is operating at fullcapacity at the point of takeoff and the Cruise engine is operating at maximum dry power with enough thrust vectored through the ventralnozzles to balance the thrust from the Lift engine. The remaining thrust is ducted through the main nozzle to acceleratethe aircraft horizontally. Figure 7.2 shows a plot of takeoff ground rolldistance as a function of the aircraft takeoff weight for all of Wh- _ cD wI- ,,,,.

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7.2 TRANSITION ANALYSIS The transition from the point of takeoff to purely wing-borne flight begins with the engines left at the same operating condition as takeoff which is with maximum vertical thrust. This configuration is held untilthe aircraft reaches a desired height. When verticalacceleration is not desired,the lift engine is throttled down and the Cruise engine starts to slowly transfer more thrustfrom the ventral nozzles to the main nozzle keeping the totalthrustbalanced about the aircraft center of gravity. The rate at which thisoccurs is such that the airplane remains level because the decrease in verticalthrust can be made to equal the increasein wing lift providing no vertical acceleration. This process is continued until the Lift engine reaches its minimum throttle setting. At this point the lift eng/nc must be shut down. During the spool.down of the Lift engine, it will still be providing some thrustbut itwill not be exacdy the desired amount so the pilot will accelerate in the vertical directionor he will have to rotatethe aircraft to a different angle of attack to alter the wing lift to compensate for the change in vertical engine thrust. This process is shown schematically in Figure 7.3. The numbers shown arc for the counter air mission, and the pilothas chosen to level off at I00 ft.altitude.

The transition from wing-borne flight to hover follows nearly the same procedure only in the opposite direction. The aircraft is brought in at a given altitudeand at approach velocity. The lift engine is started and the mar engine begins to transfera portion of the flow to the ventral nozzles. If this maneuver is done at a high angle of attack the ventralnozzles will produce drag which will significantly slow the airplane. The lift engine begins to throttle up and the Cruise engine continues to transfer more flow to the ventral nozzles to balance the force from the Lift engine. The vertical acceleration is controlled by the pilot but it is desired to keep the aircraft high enough above the ground that HGR, suckdown, and ground erosion are avoided. Once the aircraft is positioned directly above the landing site, a constant vertical acceleration of approximately 3 f't/sec is established until the aircraft touches the ground. Then the Lift engine is immediately shut down and the Cruise engine is either shut down or the ventral nozzles are closed sending allof the thrustthrough the main nozzle for ground taxiing. This is done to reduce the amount of ground erosion. A schematic of thisis shown in Figure 7.4 for the Counter Air Mission with the transition beginning at 100 ft. and the final descent to landing beginning at 50 ft.

The flightcontrol system will need alterations due to the requ/md control over the thrustvectoring. The flight control systems that ate changed for this report are: pitch and altitude hold, the bank angle control. The dynamic pressure is required for the flight control system so that the automatic flightcontrol system can determine whether to use the aerodynamic controlsof the thrustvectoring. In the block diagrams, shown in Figures 7.5 through 7.7, the "yes" by the dynamic pressure block means that the dynamic pressure is high enough to use aerodynamic controls. If the dynamic pressure is not high enough for the aerodynamic controls,thrustvectoring and the RCS will be used. The symbol _r refers to the nozzle and throttle actuation. The mason for thisis that if the nozzle deflectionsare oO

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_._ AIRPLANr' VERT. GYRO t Fimn'e7.5 PitchAttitude Hold AFC$ forthe Monarch AIRPLANE _/t T ACTuA'r.o__ REF _Ig A

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O-GYRO } Figure 7.6 Bank Angle AFCS for the Monarch I ACTUATOR H/FT ___ AIRPLANE J INO H H REF" _> --f ACTUATOR Nil E ___AIRPLANE ALTIMETER LAS Figure 7.7 Altitude Hold AFCS forthe Monarch 7.3 HOVER ANALYSIS In this section the reaction control system, suckdown predictions, and hot gas reingestion are discussed for the Monarch.

7.3.1 Reaction Control System The reaction control system maintains control about the aircraft axes in STOVL modes. It also assists the conventional control in transitional flight. Hot air is bled from the compressor of the engine and is fed to a butterfly valve which controls the flow to the four valve outlets. The butterfly valve is operated by an electromechanical actuator which is activated when the aircraft is at approach speed and below. The ducting is made from roiled and welded nickel-chromium alloy. The duct diameter was sized using Reference 7.1, the diameter varies from 4.5 inches in the fuselage to 3.5 inches at the reaction valves.

The amount of bleed air required from the engine is 2.0% mass flow of the cruise engine, this value was calculated using Reference 7.1. To calculate the amount of bleed air required for the RCS Reference 7.2 was used to find the control authority required in hover. The pitch and yaw control required for Level 1 flying qualifies in hover is .3 rad/s 2 and .5 rad/s= respectively. The pitch and yaw control required for Level 1 flying qualities in transition is .2 rad/s _ and .25 rad/s= respectively. The angular accelerations are converted into thrust by the following equation: Treq = (I x psi double dot)/1

(7.1)

With (I) being the airplane moment of inertia about the z-axis, (1) being the distance from the reaction control valve to the z-axis, and psi double dot being the yaw control required.

From the thrust it is possible to calculate the required mass flow using Reference 7.1.

The amount of bleed air for the pitch control is 1.2% mass flow of the cruise engine, and for yaw control 0.8% mass flow of the cruise engine, therefore, the total RCS bleed is 2.0%. The maximum temperature and pressure at the valves are approximately 1350 R. and 236 psi.

Roll control is provided by using variable area ventral nozzles which generate the required roll control authority for Level 1 flying conditions. The yaw control will be provided using two reaction control valves at the aft section of the fuselage. This is shown in Figure 7.8. The pitch control will be provided by using two reaction control valves at the aft section and the forward section of the fuselage. The aft section of the pitch reaction control valve is shown in Figure 7.8. The overall layout of the RCS system for the Monarch is shown in Figure 7.9.

The roll, pitch, and yaw reaction control system willbe controlled by pilot-stick movement or hover SAS. The hovcr SAS will allow the plane to remain stable throughout hover and transition.

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FOLDOUT FRAME 1. Forward RCS Pitch Valve 2. Engine Compressor Bleed 3. Butterfly Valve 4. Yaw RCS Valve 5. Aft RCS Pitch Valve H mare 7.8 RCS svs FOLDOUT FRAME layout for the Monarch' " "_ _I_ • _/I .

,' f.f Pitch Push Rod COPIED FROM RF-_RENCE 7.1 Fitnu'e 7.9 Fuselage Aft Pitch and Yaw Reaction Valve and Controls u 7.3.2 Suckdown Predictions The method used to calculate the effects of suckdown are from Reference 7.3. This method takes into account: * number of engine posts, * geometry of the aircraft, * pressure ratio at the nozzles, * in and out of ground effects, * and fountain/core effects.

The suckdown was calculated during Phase 1 of the design. The Monarch was a two post configuration in Phase 1. The suckdown predictions resulted in a 25% loss in lift versus thrust. The suckdown was assumed to be 10% for the lift engine during preliminary sizing. Therefore, the lift engine was undersized. For Phase 2 a second ventral nozzle was added so that the Monarch would become a three post configuration, m which typically reduces suckdown. The suckdown was calculated for the Monarch in Phase 2 which resulted in a suckdown of 10%. Therefore, the Monarch was changed to a three post configuration so that the lift engine did not need resizing. The comparisons between the two post and three post configurations =re shown in Figure 7.10.

w ORIGINAL PAGE IS OF POOR QUALITY The suckdown predictions are important because the hover requirement sizes the lift engine for this aircraft. Therefore, any reduction in ground effects will resalt in a lower engine weight, and eventually lead to a lower life cycle cost of the Monarch.

The equations used to calculate the suckdown for the Monarch are fi'om Roference 7.3. The resalts of the trade study between two and three post configurations is shown in Figure 7.8. Both configurations had identical geometry, the total nozzle area also remained the same for both configurations. The three post configuration needed 23% less engine thrust than the two post configuration at a height of four feet above the ground. Hgum 7.10 shows that at heights above 15 feet the three post configuration has no advantages over the two post configuration.

The reason that the three post configuration has better in-ground effects is due to the thrust "fountain core" dcvelopexi between the three nozzle posts. The "fountain core" produces lift because of the jet flow that is trapped under the fuselage due to the three separate jet flows impinging on each other. When the configuration has ordy two posts the upwash can not develop into a "core" and becomes a radial wall jet which does not produce as much lift as the three post configurations "fountain core".

-0.0 -5.C < Z w ,,,- -I0 m--r <_- -15 (./1 u- (,n Q o .J -2(3 I-- ,_J I ! I I I 0 5 I0 15 20 25 30 • rouc.oow. HEIGHT (FT) Fimn'c 7.10 Suckdown Comparison for theMonarch 7.3.3 Hot Gas Reingestion Hot Gas Reingestion is the term used to describe any flow mechanism by which hot exhaust gases from the propulsive system of an aircraft can return to the air intake of the same system. HGR is an especially important problem for STOVL aircraft operating near the ground and using propulsive Extensive theoretical research as well as full scale experiments have identified three ways in which the jet exhaust flows of a STOVL aircraft might recirculate back to the engine inlets. They are: 1) Near Field Reingestion--This is caused by the flows from separate lift jets meeting on the ground creating an upward or fountain flow which impinges on and is redirected by the aircraft undersurface. Some may travel directly on a short time scale to the engine inlets with little opportunity for mixing thereby retaining a high percentage of jet exit temperature and potentially causing severe HGR. It is shown in Figure 7.11 (Reference 7.4).

2) Mid Field reingestion or Intermediate Thrust Reverser--This is caused when some of the recirculating flow in the ground jet and the forward moving part of the fountain is blown back by headwind into the intake after some opportunity for mixing with ambient air. It is shown in Figure 7.9 (Reference 7.4).

3) Far Field Reingestion--This is caused when the ground flows navel radially outward mixing progressively with exhaust air to re,circulate into the intake on a much longer time-scale driven by the effects of buoyancy and entrainment. The rcingestion air temperature is then relatively low so Far Field Reingesdon is not usually a serious problem. It is shown in Figure 7.13 (Reference 7.4).

Fimire 7.11 Example Near Field Reingestion (_copied _om Ref. 7.4) Ficure 7.12 Example Mid Field Reingestion ('copied fi'om Ref. 7.4) Segarltid _ " ..... >>T-.7--__- _._> :: .._.<_: ._ ..::_:_.::..-.-_ :, ____-i_ ................ _"A'_ .'_',-,_t_ _..'!e = _ • t • • • All internal-combustion engines, and gas turbines in particular, are very sensitive to an increases in air intake temperatures. This arises from several causes: I) Warmer air is less dense, and the mass flow of the working fluid is therefore reduced, resulting in a loss of thrust. 2) The speed of sound in air increases with temperature, and the compressor bladeMach number at a givenrotational spcod is therefore reducc_ this reduces the compressorcapability in both non-dimensional (corrected) airflowand pressure ratio. 3) A higher air inlet temperature results in highergas temperatures throughout the engine,so that turbine temperatures become excessive; to preventthis, thrust demand must be reduced. 4) Air inlet temperana'es which change rapidly in time or space (temperature distortion) may causecompressorstall (surge) (Reference 7.4).

A major determinant of the severity of HGR is the number and location of the vertical jet exhaust nozzleson theairplane. In a nearground environment, the flow of each jetwillimpact the ground,then spread radially. If theflow of one jetmeets the flow of another jet, the flow willjoinand rise.A two jet configuration willresult in a long wall of upward flow being generated between the two jets.A three-jet configuration will produce a concentrated fountain at thepointwhere the flow of allthree jets combine.

There will also be threc walls extending from this fountain where two of the jetscombine.

A four jet configuration (like theHarrier) will produce a more concentrated fountain at the center of thefourjets and four wallswill extend from it. Itis the upward airflow that willreachtheinlets and cause Near Field Reingcstion so regardless of the number of nozzles, itisdesiredto keep the fountain and wall airflows away from theinlet area.

7.14 isa top-view of theMonarch showing the location of all nozzles and inlets as wellas the fountain thatis created by the engine flow. Noticethatthereisnot a wall of airflow under the fuselage at the location of theinlets.

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There has been extensive research into many methods of alleviating HGR by making slight modifications to an airplane configuration. If HGR does become a major problem for the Monarch, then one or more of these modifications should be made. The following list shows some of these ideas.

1)

Attempt to deflect the Lift/Cruise nozzles outboard ff there is not a significant loss in thrust. A slight deflection, as shown in Figure 7.15, wiU sufficiendy direct the flow away from the airplane. Studies have shown that this may alleviate the ITR enough that the loss of thrust due to the nozzle angle is more than compensated by the improve engine performance. (Reference 7.5)

2)

Place deflector shields near the nozzles to direct the flow away from the inlets. It may be possible to integrate current doors to the landing gear and the missile bay to also act as this type of shield, or it may be necessary to make separate shields that retract into the fuselage. Figure 7.16 demonstrates using a door that covers the Lift engine nozzle.

3)

Create an "air curtain" around the inlets by ducting compressor air from the Lift/Cruise engine out of the fuselage near the inlets. This air flow will entrain and remove the hot gases that would otherwise enter the inlets. According to Reference 7.Y, approximately 2 % of the engine air flow is necessary to create this type of curtain. An approximate location as well as a schematic of the air flow is shown in Figure 7.17.

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7.4 PILOT WORKLOAD ANALYSIS Control requirements and pilot workload for STOVL aircraft are higher than that of conventional aircraft. The STOVL aircraft may be r_luired to operate from conventional airfields, austere sites, and aircraft carriers. The capability for hover and low-speed flight and for rapidlytransitioning between wing-borne and propulsion-borne flights permits the STOVL aircraft to operate into confined spaces associated with austere sites. These operations enforce precision of control of position, velocity, and attitude; such requirements exceed those imposed on conventional aircraft (Reference 7.6).

A major technological challenge to routine vertical flight operations of this class of aircraft in adverse weather and low-visibility conditions stems from the complex interaction of kinematics, aerodynamics, and propulsiveforces and moments during transition as reflected in poor flying qualities as well as from limited control authorities. The availability of digital fly-by-wire controls makes it feasible to reduce the amount of pilot workload during takeoff, transition, and hover. To also help in reducing the pilot workload the number of control sticks will be reduced from three (Harrier AV-SB) to two. The digital fly-by-wire controls and the advancements made in flight control software will allow for reduced pilot workload.

The cockpit controls and displays for the Monarch are adapted from Reference 7.6.

The cockpit controls and displays for transition is shown in Figures 7.18 and 7.19. The situation/director display ('Figure 7.18) is a three-cue compensary flight director supplerncntedby situation informationpresented in both analog and digital format. The flightpath pursuit/situation display (Figure 7.19) projects a lead aircraft that is following the desired flightproRle. The cockpit controlsand displays for hover are shown in 7.20.

The HUD format in transition and hover is shown in Figures 7.21 and 7.22, respectively.

The workload for the pilot at takeoff is reduced because the flight control software performs the nozzle and control surface deflections to minimize the takeoff distance. The methodology of the flight control system for takeoffis discussed in Section 7.1.

For landing the pilot will bring the aircraft to approach speed and at that time the pilot will have the option to select "landing". If the pilot selects landing the HUD will switch over to situation/director display and the cockpit controls will switch over for transition (Figure 7.18). The pilot will then be given the option to select the landing location with the Forward Looking Infi'a Red (FLIR). Once the landing locationis selected the HLrD will display the pursuit/situation display and also the cockpit controlswill switch to the flightpath-centered pursuit (Figure 7.19). The pursuit/situation HUD will allow the pilotto follow the ghost plane and also the landing locationwill be displayed on the HUD.

When the aircraft gets within hover range the hover HUD will be displayed and the flight controls will switch to the hover mode (Figure 7.20). The hover HUE) will allow the pilot to see the desired hover point along with the other important information as shown in 7.22.

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Cockpit Controls and Displays for Hover (Copied from I_ference 7.6) Fixate 7.20 DISTANCE TO STATION KEEPING POINT %RPM ""'_ ILS 7S _ THRUST VECTOR ENGINE ._ ANGLE ----8 _" 1.7 AIRCRAFT PITCH __ PITCH REFERENCE LADDER •-== 4 HORIZON WITH 60 1,jJ HEADING SCALE i l I I I w LATERAL ACCELERATION -3 dq GLIDESLOPE I" REFERENCE ALTITUDE AIRCRAFT GHOST w Figtn'e 7.21 HUD Format in Transition (Copied from Reference 7.6) L ABOVE DECK 50 40 - - AIRSPEED VERTICAL VELOC_v HORIZONTAL SPEED wJ.t. SHiP DECK ALLOWABLI VERTICAL VERTICAL AIRCRAFT VELOCITY REFERENCE (VERTICAL (TRIDENT) PERSPECTIVE) Figme 7.22 HUD Format in Hover (Copied fi'om Reference 7.63 REFERENCES FOR CHAPTER 7 7.1 McDonnell Douglas Corp. Report #MDC A7726 7.2 AGARD Report No. 710., Special Course on V/STOL Aerodynamics, 1984.

7.3 Henderson, C.; Clark, J.; and Waiters, M.; "V/STOL Aerodynamics and Stability and Control Manual", NAVAIRDEVCEN-80017-60, January 1980.

7.4 Penrose, C.J., '"1"he Scaling of Model Test Results to Predict Intake Hot Gas Reingestion for STOVL Aircraft with Augmented Vectored Thrust Engines".

7.5 Ryan, Patrick E., Helm, Richard J., and Cosgrove, Wayne J., "A Generalized Experimental Investigation of Hot Gas Recirculation and Ingestion for Jet VTOL Aircraft", NASA CR-1147, September 1968.

7.6 Foster, John D., and Schroeder, Jeffrey A., "integrated Control and Display Research for Transition and Vertical Flight on the NASA V/STOL Research Aircraft", NASA TM 100029, October 1987.

8. PERFORMANCE DATA AND MISSION CAPABILITY The purpose of this chapter is to present the performance dam and mission m capability of the Monarch aircraft. The drag characteristics of the aircraft is summarized and shown in Section 8.1. Section 8.2 presents the performance data and Section 8.3 presents the mission capability. The spreadsheetsused to calculate the performance data and mission capability are shown in Appendix 2.

8.1 SUMMARY OF DRAG CHARACTERISTICS The drag polars of the aircraft were calculated and are fully documented in Reference 8. I. The drag polars were adjustedto account for trim drag in Reference 8.2.

The Monarch drag polars are shown in Table 8.1. The validity of the drag calculations is shown using Figure 8.1,where the skin friction coefficient of the Monarch is compared to similar aircraft. The wave drag for the configuration was calculated using the method of Reference 8.3 and actual data for the Grumman 1=-14 and the General Dynamics F-16 taken from Reference 8.4. The Monarch wave drag is shown in Figure 8.2.

Table 8.1 Monarch Drag Polars I-I(fl) M Zero Lift Drag. C_o Induced Drag Factor. 1/(PI*A*e) 0 0.20 0.02198 0.1091 100 0.85 0.02096 0.1022 10,000 0.90 0.02281 0.1002 15,000 0.90 0.02410 0.1003 30,000 0.90 0.02750 0.1103 30,000 1.20 0.04157 0.1006 30,000 1.60 0.04038 0.1008 40,000 0.80 0.02387 0.1103

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8.2PERFORMANCE DATA The following performance data ispresented: * point performance verification * sustained turnrateand load factor * specific excessenergy * maximum ferryrange The Monarch also compared to operational fighters in the United States and SovietUnion to show itsvalidity as a designand its combat effectiveness against theseaircraft.

Point Performance Verification Table 8.2 shows the point performancerequirements from themission specification and thevalues calculated for theMonarch. Note: Allperformancedatapresented arefor a combat weightof 26,192 Ibswhich includes 50% fuel, two short range missiles, and half the ammunition for the cannon.

Table 8.2 Point Performance Verification for the Monarch

M.ontmb..Y_al_

P¢fforr_ance Requirement

Time to Climb 40k in 2 minutes 1.75 rain Ig Specific ExcessEnergy 500 ft/sec 505 f-t/see (2A) 30k 0.9M 1,000 ft/sec 920 ft/sec (2B) 10K 0.9M Sustained Turn Rate l (3A) 0.SM/15k ft 15 deg/sec 15 deg/sec (3B) 0.9M/30k ft 9 deg/sec 10 deg/sec (3C) 1.2M/30k ft 8 deg/sec 9.9 deg/sec (3D) 0.9M/15k ft 6.5 g 7.75 g (3E) 1.6M/30k ft 4.5 g 8.70 g Acceleration (4A) 30k ft 0.9M to 1.6M 70 sec 47.3 sec (4B) 0.5M to 1.4M 80 sec 62.1 sec (4C) 1Ok ft 0.3M to 0.9M 22 sec 18.4 sec Landing Distance (ground roll) Without Chute 2,200 ft 2,100 ft The Monarch meets all it required performance except for the 1000 ft/sec specific excess energy requirement. The improved performance of this a_raft as compared to the Phase I study isdue to the upsizing of the engineof the Monarch which was done in the Phase IIiteration.

Sustained Turn Rate and Load Fire)or The sustained ttu'n rate and load factor were calculated for the Monarch and are shown in Figures 8.3 and 8.4, respectively. The Monarch is capable of a sustained mm rate of 21 dee/see at low level and can sustain 6 deg/sec at altitudes as high as 45,000 ft.

A sustained load factor of 9 g's is maintained for much of the low altitude and high Mach number flight envelope. A 3 g sustained load factor is achievable at altitudes up to 50,000 ft.

The Monarch's turn performance at 15,000 ft is shown in Figure 8.5. This "dog house" plot shows the relationship between turn rate, load factor, turn radius, and Math number. This plot shows that the Monarch can sustain high rates of turn over the operating Math number range due to its high thrust engine. The maximum sustained turn rate at 15,000 ft for the aircraft is 16.9 dee/see (thrust limited) and the maximum instantaneous turn rate is 17.3 dee/see (lift limited).

Specific Excess Ener_ The l g specific excess energy for the flight envelope was calculated for the Monarch and is shown in Figure 8.6. The Monarch has a 1,000 ft/see specific excess w energy at high subsonic Mach numbers at altitudes below 10,000 ft. A specific excess energy of 600 ft/see is achievable over a wide part of the high Mach number flight envelope.

Maximum Ferry. Range The maximum ferry range calculations arc plotted in Figure 8.7. The maximum range of the aircraft is 1662 nm at 45,000 ft and M - 0.9. Range credit for climb was included in the calculations, as well as fuel use for climb, descent, and takeoff. This amount of range is feasible with fuel tanks fitted into the internal weapons bay volume.

The aircraft uses two cylindrical tanks for this application.

The takeoff maximum thrust and combat weight versus Mach number are shown for the Monarch and several other fighter, respectively, in Figures 8.8 and 8.9. The Monarch fits into the trend of these other aircraft within reason.

As a measure of the Monarch's combat effectiveness, its turn rate and agility potential are compared to several other fighters. These plots arc shown in Figures 8.10 and 8.11, respectively.The sustainedturn rate (at 15,000 ft and Mach 0.9) of the Monarch exceeds the instantaneousturn rateof the Mig-21, Mig-23, and the F-15. The Monarch and the F-16 have comparable turn capabilities at thisMach and altitude.The agility potential shown is one of the only static agility metric available and is defined as: AgilityPotential = (T3_/WTo)/(WomJS) The Monarch compares favorable to the F-14 and F-16, but falls short of the agility potential of the F-15. A lower wing loading for the Monarch would improve this ability, but would then make the aircraft less comfortable on a bombing mission.

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8O 4.O P'--__. _'--15 _'--16 AI_C_ Fieur_ 8.11 A_HIy Po_nH_l Comparison 8.3 MISSION CAPABILITY The mission capability of the Monarch is measured by fast verifying the design missions and second, taking the aircraft through typical fighter/attack missions to determine the aircraft's capability as a multi-role fighter. Tables 8.3 and 8.4, respectively, show the counter air mission and battlefield air interdiction missions fuel usage for the design mission. The supersonics (acceleration to and sustaining supersonic flight) of the counter airmission and the low level dash of the battlefield air interdiction mission dominate the aircraft fuel usage.

u Table 8.3 Counter Air Mission Fuel Burn Summary.

Phase Fuel Burn 314 lbs 1. Engine Start/Warm Up 2. Taxi 279 lbs 3. Short Takeoff 360 lbs 313 lbs 4. Acceleration to Climb Speed 5. Climb 485 Ibs 6. Subsonic Cruise - 100 nm 531 lbs 620 lbs 7. Acceleration to Supersonic Oxtise 1334 lbs 8. Supersonic Cruise - 50 nm 9. Combat 1728 lbs 1325 lbs 10. Supersonic Cruise - 50 nm 11. Subsonic Cruise - 100 nm 571 lbs 12. Hover 227 lbs 114 lbs 13. Landing 14. Reserves 432 lbs CA Mission Fuel Burn = 8634 lbs u Table 8.4 Battlefield Air Inteniiction Mission Fuel Bum Summary_

Phase Furl Burn

327 lbs 1. Engine Start/Warm Up 2. Taxi 307 lbs 3. Short Takeoff 376 lbs 308 lbs 4. Acceleration to Climb Speed 5. Climb 538 lbs 6. Subsonic Cruise - 200 nm 1331 lbs 7. Sea Level Dash In - 80 nm 1204 lbs 8. Strafe Run 864 lbs 9. Sea Level Dash Out - 80 nm 1110 lbs 10. Climb 326 Ibs 11. Subsonic Cruise - 200 nm 1124 Ibs 12. Hover 246 lbs 121 lbs 13. Landing 14. Reserves 432 lbs Battlefield Air Interdiction Mission = 8614 Ibs Typical NATO fighter/attack mission profiles were obtained from Reference 8.5.

The missions arc: Figure 8.12 Mass Intercept Figure 8.13 Transport/Hellcopter Intercept Figure 8.14 AWACS/I-Iigh Value Asset Protection w Figure 8.15 Two Stage Mission The figures show the Monarch's range and speed capability in these missions. The high value asset protectionmission and the two stage mission offer unique advantages for a STOVL type aircraft. As shown in the profiles, a STOVL aircraft can operate from dispersed bases and thus save fuel and cut down on response time.

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I r REFERENCES FOR CHAPTER 8 8.1 Cox, Brian, et. al., Preliminary. Dcsi_ of a Su_rsonic STOVL Aircraft Using a Lift + Lift/Cruise En_ne Cycle, University of Kansas, AE 621, December 1989.

8.2 Cox, Brian, et.al., Preliminary Stability and Control for the Lift ConfimLration.

University of Kansas, AE 622, March, 1990.

8.3 Raymer, Daniel P., Ah'craft Dcsi_: A Oasnceptual Approach, American Institute of Aeronautics and Astronautics, Washington, D.C., 1989.

8.4 Hammond, David, "CAM Aero Report", AFWAIJTXAD, 2 November 1988.

8.5 Wright Research and Development Center, Technology Assessment Division In-House Design Mission Profiles, Co_dinator: 2nd LL Gerald Swift, January 1990.

8.6 Northrop Lecture Series to NASA Ames, NASA Ames Research Center, Moffet Field, CA, Summer 1989.

9. STABILITY AND CONTROl, The purpose of this chapter is to document the results of the stability and control analysis for the Monarch fighter. The following topics are covered in this chapter: 9.1 Flight Conditions 9.2 Trim Diagrams 9.3 Stability and Control Derivatives 9.4 Dynamic Stability and Con_ol Analysis 9.4.1 Longitudinal 9.4.2 Lateral 9.4.3 Directional 9.5 Roll Performance 9.6 Inertia Coupling 9.7 Spin Departure 9.8 Low Level Ride Qualities 9.9 Vertical Tail/Rudder Removal Study 9.1 FLIGHT CONDITIONS This section presents the selection of eight flight conditions which are representative of the flight envelope of the Monarch. A description and list of the parameters of each flight condition is also given.

Eight flight conditions were chosen to represent the flight envelope of the Monarch fighter. They were chosen from the Counter-Air (CA) and Battlefield Air Interdiction ('BAD Mission profiles as depicted in Figures 9.1 and 9.2, respectively.

A description of the flight conditions follows: FC 1: CA, Phase 3, Takeoff/Hover/Transition phase.

FC 2: BAI #1, Phase 6, Low altitude, high subsonicdash out to ordnance drop.

FC 3: CA, Phase 8, Subsonic performance point.

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FC 5: CA, Phase 5, Subsonic maneuver, load factor = 6.0.

FC 6: CA, Phase 9, Supersonic performance point.

FC 7: CA, Phase 7, I-ligh altitude, supersonic cruise.

FC 8: BAI #2, Phase 5, High altitude subsoniccruise.

Table 9.1 summarizes theparameters of each flight condition.

Table 9.1 Flight Conditions for the Monarch Fighter FLight Condition Altitude Mach Number I,d_L,F_ggl_ 1 0 ft 0.20 1.0 2 100 ft 0.85 1.0 3 I0,000 ft 0.90 1.0 4 15,000 ft 0.90 4.25 5 30,000 ft 0.90 6.0 6 30,000 ft 1.20 l.O 7 30,000 ft 1.60 1.0 8 40,000 ft 0.80 1.0 9.2 TRIM DIAGRAMS This section presents thetrimdiagrams for theMonarch fighter. The method of Reference 9.1was used in constructing thetrimdiagrams. Detailed calculations of thetrim data are documented in Reference 9.2.

m The airplane lift versus angle of attack curve and airplane lift versus pitching moment curve were calculated according to the methods of Reference 9.1. The eta'yes were constructed forhorizontal tail deflections ranging from -30 degrees to +30 degrees in ten degreeincrements.

The forwardand aft c.g.travel linesand the horizontal tail stall loci form the i boundariesof theaim triangle. Within the bounds of the trimtriangle, the horizontal stabilator deflections necessary to aim the airplane for a range of lift coefficients are determined. For each flight condition the lift coefficient was determined, knowing the airplane weight, load factor, altitude and velocity. From the respective trim diagram it was determinedifthe airplane could be trimmed and, ff so,what stabilator deflection was require& At the end of Phase I design, theMonarch was designed with a canardfor longitudinal control. The detailed stability analysis required for the developmentof the trimdiagramsrevealed thatthe canard designhad an unacceptable margin of longitudinal instability. This ledto theremoval of the canardand theincorporation of a conventional tail aftstabilator intothe designof theMonarch.

Originally a symmetricairfoil was selected for the horizontal stabflator design. The symmetric airfoil displayed a low stall angleof attack and made the airplane untrimmable in all flight conditions. A cambered 6% thick airfoil was incorporated to improve the tail stall characteristics. In addition, a full span fixed slat, similar to the stabilator design on the McDonnell F-4E Phantom II, is used. These changes provided adequate longitudinal control power throughout the e.g. ranges of all flight conditions. According to criteria found in Reference 9.3, the drag divergence Mach Number of the horizontal stabilator and the wing were determined. From this it was determined that the drag divergence Mach Number of the stabilator was higher than that of the wing. Therefore the horizontal stabilator will retain control power at high subsonic Mach Number when the flow over the wing becomes supersonic.

The trimdiagramsfortheMonarch fighter are shown in Figures 9.3 through9.10 for the eight flight conditions. Where the center of gravity limits cut intothetrimmable range of the aircraft, the fuelmanagement system willkeep the center of gravity from moving intothese areas. This willkeep the aircraft prevent from moving intountrimmable flight conditions.

The trimdiagram for flight condition I (Figure 9.3) reflects the lift increments and corresponding pitching moments fora 40 degreetrailing edge plain flapand 20 degree leading edge slat. The thrust from the lift engine and the main engine are balanced to augment the aerodynamic lift duringtakeoff as described in Chapter 1I.

From thetrimdiagrams it was determinedthatthe Monarch can be trimmed with reasonable stabilator deflections forall flight conditions. Table 9.2 lists the lift coefficient and the required stabilator deflection to trim foreach flight condition.

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"1_ './.N_IOI_300 J...-II'l w 17,,4 Table 9.2 Required Stabil_lor Deflection to Trim Stabilator Deflection. de_ FC Lift Coefficient 1 1.495 0.0 2 0.076 -1.5 3 .083 -2.0 4 .661 3.7 5 .968 6.0 6 .111 -3.9 7 .069 -3.0 8 .489 3.0 9,3 STABILITY AND CONTROL DERIVATIVES The stability and control derivatives for the Monarch fighter arc presented in this section. The methods of References 9.1 and 9.4 were used to compute the derivatives for the eight flight conditions. Reference 9.2 documents the detailed calculations of the stability and control derivatives for the Monarch fighter.

The longitudinal, lateral-directional and thrust derivatives of the Monarch fighter for the eight flight conditions are presented in Tables 9.3 through 9.10. Also presented in these tables are the geometric and flight condition parameters required for the calculation of the dimensional derivatives.

The thrust derivatives were calculated with data obtained from the Pratt & Whitney engine deck (Reference 9.5).

Due to the fact that the rudder had been eliminated (Section 9.8), directional control was achieved using the 2-D main vectoring nozzle. Because the vanes for vectoring the thrust directionally are comparable to a control surface (6 sq. ft.), aerodynamic control derivatives for the vanes were calculated along with the control derivatives due to the thrust vectoring.

The longitudinal and lateral-directional derivatives of the Monarch fighter were compared to data of Reference 9.6. Reference 9.6 is a graphical presentation of the stability and control derivatives for supersonic fighters as a functionof Mach Number.

Figures 9.11 through 9.16 are copies of selecteddam from Reference 9.6 with the values for the Monarch fighter included. The values for the Monarch are illustrated with a circled dots in these figures.

9.4 DYNAMIC STABILITY C_DNTROL AND ANALYSIS Due to time constraints, three flight conditions were chosen for dynamic stability and control analysis. These conditions were chosen to cover the least similar flight regimes. Flight condition 2 was chosen to represent a high speed, low altitude ground attack phase. Flight condition 4 is representative of air-to-air combat at a typical Table 9.3 Stabili_ and Conm:_l Derivatives for Fli2ht Condition 1 Geometric and Flight Parameters: S = 347.9 (sq ft) I xx S - 13785 (slug* ft^2) b = 33.67 (ft) I_yy_S ffi 90780 (slug*ft^2) c bar = 12 (ft) I zz S = 82563 (slug*ft^2) mass ffi 954.8 (slugs) I xz S = -13839 (slug*ft^2) Longitudinal: Lateral-Directional: C D 1 ffi 0.2660 -0.4630

C_y_B ffi

C D u = 0.0000 0.0130

C_y-2 ffi

C D a= 1.1060 0.2580 C_y_r = CD ih= 0.0000 0.0000 C_y_dA = C T x 1 = 0.2660 w 0.0183 (aero) C_y_dR = C T xu ffi -0.5091 CydR= 0.3043 (thrust) C M T 1 = 0.0089 -0.2394 C M T u = -0.0177 -0.2155 0__1_2 = C M T a ffi 0.1024 C1 rffi 0.3810 C L 1 = 1.4950 CY_A= 0.4150 C L u ffi 0.0440 C--l--dR -- -0.0029 (aero) C L a ffi 3.7470 C-l-dR = -0.0497 (thrust) C L adot = 0.5182 CnB= -0.0034 5.9700 CnTB= 0.0259 C_L_q = C L ih ffi 0.3970 -0.0900 C_n_p = CMI = 0.0647 Cnr= -0.2980 CMu= 0.0000 CndA= -0.1630 CMaffi 0.1490 CndR= -0.0111 (aero) C M adot = -0.5959 CndR-- -0.1904 (thrust) -4.2060 C_M_q = CM ih ffi -0.5280 Table 9.4 Stability and Control EMrivativcs for Flight Condition 9, Geometricand Flight Parameters: S = 347.9 (sq ft) I xx S = 15370 (slug*ft^2) b = 33.67 (ft) I yy_S = 88824 (slug*ft^2) c bar = 12 (ft) I zz S = 86851 (slug*ft^2) mass = 881.8 (slugs) I xz S = 1035 (slug*ft^2) Longitudinal: Lateral-Directional: -0.5030 C_D_I = 0.0241 C_y_B = C D u = 0.0850 -0.1150 __ C y_p= C D a = 0.0621 C--v r 0.3110 C D Th = 0.0102 C 0.0000 C T x 1 = 0.0241 C_y_dR = 0.0183 (aero) C T x u = -0.0467 C_y_dR = 0.0239 (thrust)

c MT 1 -- 0.0008 c 1 B =

-0.0708 -0.2840

c M T u = -o.oo16 cillp

C M T a = 0.0217 C 1 r = 0.1650 C L i = 0.0760 C _ _IA = 0.2520 C L u = 0.0820 C--l--dR = -0.0006 (aero) C L a = 4.4380 C--l--dR = -0.0008 (thrust) C L adot = 0.6249 _ _ B = 0.0112 C_L_q = 9.2080 C n T B = 0.0056 0.0680 C L ih = 0.5040 C_n_p = C M i - 0.0076 C n r = -0.3200 C M u = -0.0057 C n dA = -0.0050 C M a = 0.4420 C n dR -0.0115 (aero) C_M_adot = -0. 6838 C n dR = -0.0150 (thrust) C M q = -6. 9500 -- -- C M _h = -0.6700 Table 9.5 Stability and Control I)crivadvesfor Flight Condition 3 Geometric and Flight Parameters: S = 347.9 (sq ft) I xx S = 9904 (slug*ft^2) 87959 (slug*ft^2) b = 33.67 (ft) I_yy_S = c bar = 12 (ft) I zz S = 81040 (slug*ft^2) mass = 744.7 (slugs) I xz S = 820 (slug*ft^2) Longitudinal: Lateral-Directional: CD 1 = 0.0246 -0.6010 C__y__B = C D u = 0.1260 -0.1370 C__y...p= CD a = 0.0751 0.3720 C_y_r = C DTh = 0.0147 0.0000 C_y_dA = C T x 1 = 0.0246 0.0183 (aero) C_y_dR = C T x u = -0.0429 CydR= 0.0533 (thrust) CMT I- 0.0009 _YB -0.0832 CMTu= -0.0015 -0.2950 CMT a= 0.0307 C 1 r = 0.1780 C L 1 = 0.0830 cY A = 0.5110 C L u = 0.1700 CldK-- -0.0006 (aero) C L a = 4.5770 CldR= -0.0018 (thrust) C L adot = 0.6393 C n B = 0.0665 m -- 10.5510 C n T B = 0.OO78 C_L_q = C L ih = 0.5190 0.0790 C_n_p = CMI= 0.0090 Cn r = -0.3140 CMu= -0.0111 CndA= -0.0110 CMa= 0.4920 CndR= -0.0114 (aero) C M adot = -0.6952 CndR= -0.0322 (thrust) -8.3400 c M_q CM lh _ -0.6900 u.

m Table 9.6 Stability_ and Control Derivatives for Flight Condition 4 Geometric and Flight Parameters: 10631 (slug*ft^2) S = 347.9 (sq ft) I xx S 87959 (slug*ft^2) b = 33.67 (ft) I_yy_S 80313 (slug*ft^2) c bar = 12 (ft) I zz S = -7202 (slug*ft^2) mass = 744.7 (slugs) I--xz--S = Lateral-Directional: Longitudinal: -0.6010 CD I = 0.0688 C_y._B = -0.0900 w C D u = 0.1260 C_y_p - 0.3860 CD am 0.6064 C_y_r - 0.0000 C D ih = 0.0272 C_y_dA = 0.0183 (aero) C T x 1 = 0.0688 C_y_dR - w 0.0763 (thrust) C T xu= -0.1220 -0.1600 CMT 1 = 0.0024 -0.2950 -0.0042

C M T u = c-x-p =

w C i r= 0.1650 0.0383 CMT a= = 5.1100 0.6610 C L 1 = CIdR- C--L--u - 0.6500 -0.0019 (aero) C--L--a = 4.5770 CIdR-- -0.008 (thrust) 0.0802 0.6393 C n B = C L adot = 0.0098 10.5510 Cn TB= --_ L q 0.0200 C_lh= 0.5190 C_n_p = Cn r = -0.3670 CM 1 = 0.0711 CE A- -0.0080 CMu= -0.0879 CndR= -0.0010 (aero) CMa= 0.4920 CndR= -0.0469 (thrust) C M adot = -0.6952 -8.3140 C_M_q = C M ih = -0.6900 Table 9.7 Stabili_ and Control Derivatives for Flight Condition 5 Geometric and Flight Parameters: S = 347.9 (sq ft) I xx S = 9545 (slug*ft^2) b = 33.67 (ft) I--yy--S = 90165 (slug*ft^2) c bar m 12 (ft) I zz S = 987 (slug*ft^2) mass - 876.8 (slugs) I xz S = 3283 (slug*ft^2) Longitudinal: Lateral-Directional: CD 1 = 0.1292 C_y_B = -0.6010 C D u = 0.1260 C_y_p = -0.0650 CD a= 0.9840 C_y_r = 0.3910 C D ih = 0.0441 C_y_dA = 0.0000 C T x 1 = 0.1292 C_y_dR = 0.0183 (aero) C T x u = -0.1867 CMT 1 = C y_dR i B = 0.1385 (thrust) 0.0045 -0.2050 CM T u = -0.0065

c21 = -02950

C M T a = 0.0728 C i r = 0.1200 C L 1 = 0.9680 C T dA = 0.5110 C LU= 0.O450 C i dR = -0.0027 (aero) w C L a-- 4.5770 C--l--dR = -0.0204 (thrust) C L adot = 0.6598 n B = 0. 0941 10.5510 C_L_q = C n T B = 0.0861 CL ih= 0.5190 C_n_p = -0.0140 CM 1 = 0.0698 C n r = -0.4000 CMu= -0.1287 C n dA = -0.1290 C M a = 0.3300 C n dR = -0.0111 (aero) C M adot = -0.7430 C_n_dR - -0.0858 (thrust) -8.3140 -- _ M q C_ lh == -0.6900 w Table 9.8 Stability. and Control I_rivatives for Flight Condition 6 Geometric and Flight Parameters: S = 347.9 (sq ft) I xx S = 6694 (slug*ft^2) b = 33.67 (ft) I_yy_S = 88352 (slug*ft^2) c bar = 12 (ft) I zz S = 78707 (slug*ft^2) mass = 761.9 (slugs) I xz S = 486 (slug*ft^2) Longitudinal: Lateral-Directional: CD 1 = 0.0416

Cj_B = -0.6260

CD u= -0.1200 C_y_p = -0.1390 CD a= 0.0817 C_y_r = 0.4130 C _Th = 0.0181 C_y_dA = 0.0000 C T x 1 = 0.0416 C_y_dR = 0.0183 (aero) C T x u = -0.0303 C YYdR=B 0.0453 (thrust) CMT 1 = 0.0014 = -0.0885 C M T u = -0.0011

C=l=p = -o.335o

C M T a = 0.0524 C 1 r = 0.1630 C L 1 = 0.Iii0 C [dA = 0.1630 CLu= -0.0500 C 1 dR = -0.0007 (aero) CL a= 3.7980 C 1 dR = -0.0018 (thrust) C L adot = 0.3244 C n B = 0.1048 10.1420 C_L_q ffi C n T B = 0.0133 C L ih = 0.3990 C_n_p = 0.0770 C M 1 = -0.0011 C n r = -0.4160 CMu= -0.0154 C n dA = -0.0047 CMa= -0.0390 C n dR = -0.0114 (aero) C M adot = -0.3764 C_n_dR = -0.0282 (thrust) -6.8550 C = -0.4510 Table 9.9 _Stability. and Control Derivatives for Fli_ht Condition 7 Geometric and Flight Parameters: 10152 (slug*ft^2) S = 347.9 (sq ft) I xx S - 8967 (slug*ft^2) b = 33.67 (ft) I_yy_S = c bar = 12 (ft) I zz S = 83546 (slug*ft^2) 410 (slug*ft^2) mass = 845.9 (slugs) I--_xz--S = Lateral-Directional: Longitudinal: C D 1 = 0.0408 C_y_B - -0.4250 0.0000 C D U = C_y_p = -0.0940 0.0300 C D a = C y r = 0.2800 C 5Th = 0.0084

c____ = o.oooo

O.0408 C T x 1 = C_y_dR = 0.0183 (aero) u -0.0420 C T xu = C_y_dR = 0. 0428 (thrust) 0.0014 C 1 B = -0.0590 C M T 1 = -0.0015 CM T u =

-0.2750

0.0329 C 1 r = 0.1530 CMT a= C T _A = 0.0970 C L 1 = 0.0690 C--l--dR = -0.0007 (aero) C L u - -0.1550 2.3780 C--l--dR = -0.0017 (thrust) C L a = C L _d_t = 0.2569 C n B = -0.0083 5.0180 C n T B = 0.0080 --_ L q C_xh_ 0.4440 C_n p = 0.0560 C n r = -0.4320

_i = -0.0033

m

C--M--u c _A = -0.0018

0.0000 C--M--a -0.1130 C--n--dR = -0.0118 (aero)

c M E_t -0.3034 C--n-dR = -0.0270 (thrust)

-3.6780

- g M._q

Cg_h_ -0.5110 m = = Table 9.10 Smbilit'y and Control Derivatives for Flight Condition 8 Geometric and Flight Parameters: S = 347.9 (sq ft) I xx S = 13169 (slug*ft^2) b = 33.67 (ft) I_yy_S = 88119 (slug*ft^2) c bar = 12 (ft) I zz S = 83378 (slug*ft^2) mass = 929.1 (slugs) I xz S = -5241 (slug*ft^2) Longitudinal: Lateral-Directional: CD 1 = 0.5350 -0.4910 C_yB = CDu= 0.0510 -0.0840 C__y....p ,- CD a= 0.4770 0.3120 C_y_r = C D ih= 0.0189 0.0000 C_y_dA -- C T x 1 = 0.5350 0.0183 (aero) C_y_dR - C T x u = -0.0188 0.0536 (thrust) C_y_dR = CMT 1 = 0.0018 C 1 B= -0.1263 CMTu= -0.0007 -0.2690

cD-p =

CMT a= 0.1282 C 1 r= 0.2100 C L 1 z 0.4890 CIdA= 0.3990 C L u = 0.0200 CIdR= -0.0016 (aero) C L a = 4.3190 C 1 dR= -0.0046 (thrust) C L adot = 0.6059 CnB= 0.0436 8.4730 Cn TB= C_L_q = 0.0328 C L ih = 0.4850 0.0300 C n_p = CM 1= 0.0452 Cn r = -0.2950 CMu= -0.0180 CndA- -0.0431 CMa= 0.3990 CndR- -0.0116 (aero) C M _d_t = -0.6672 CndR-- -0.0334 (thrust) -6.2550 --_ M q C_ lh == -0.6460 S = 347.9 so ft b = 33.67 ft c_bar = 12.0 ft u m Fi_c 9.11 Variation of C L___cith Mach Number S = 347.9 sq ft b = 33.67 ft char = 12.0 f_ llotlce the _ther mall v&lues elh_.bltKl by the tallleH ¢ont:llurstlon.

. The estimated rlatge ot vslues of Cot for present ,,,d near-tuture Jet tighter typo &lrcrsft is troa-20 to O.

I I I I I • I Copied from l_fcrence 9.6 _---14.

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Variation of (2 M q with Mach Number Fi_._ 9,12 ORIGINAL PAGE IS OF POOR QUALITY S = 347.9 sq b ffi 33.67 ft c_bar = 12.0 J i I J I a , l _'_e ooaclude s that decreuinl C_B /a_roves the OTtTLLI fllcbt behavior.

(DrLke. H.M.. "_he I_fect of Lstertl Ares on the Lsteral 8tsbllLt.y ud O;)r_t.z'oJ. OAar_'t, eri.st.l_.l Of Lm _t'_lll_e U _JCemtned _ '_Stll Of • Model In the Langley PTee-PIAIht Tunnel. " NACA Advs_ce Kes_TtCtod Report..

AM L,SL05, Lanjle_v llmorl-] kerm&utLcal Laborstory. L_sle7 P/eld. Vs..

PebruLry 1946. ) _ere Is no I_u'_t oorFelatton betmn Cr& vt_uel _ wU_I )lutom t_pe. lheestlasted rsnse of values of CwD tot present end nest-rue-re Jet flsh_Jr type aircraft is from -.1 to -1.5.

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Fi_r_ 9.13 Variation of C y B with Mach Number OP,_C_.H,GL PAGE fS OF POOR QUALiTy

I i

$ = 347.9 sq ft b = 33.67 ft c_bar = 12.0 ft I ! I The esttnated ranl_e of values of C_81 for present tnd near=future Jet fijhter _ype alrcrsft is fz_ 0 to..5.

roB. 48" Copi_l from Reference 9.6 m--, 40" - _ --.32, c ": --2, -. 16.

w i m-,08_ mo • 4 .8 1.2 J Lq; 2.0 2.4 _.| hch Nmber

Ill

Fi_xtre9.14 Variation of C y _SR with Math Number m lind lear*future Jet DT@8_t 7he estLsated rsn|e of values of C_jfor _lshter t,Tpe aircraft Ls fral 0 to. 40.

-- ..35--

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Ii

Variadon of C n B with Mach Number

III

S = 347.9 sq ft b = 33.67 ft c_baz = 12.0 ft Por all configurations, rudder effecUveaess decreues Ur_ot].v An the the transonic regloa. ?he eatAaated raase of val-u of Cm# a for prueat and near-future let fXl_ter t_l)e sArcrttt ts free O to -. 15.

Copied from Reference 9.6 --. X21 -. 10 P-80A

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I J

Fimm: 9.16 Variation of C n _b'R with Mach N-m_'r engagement altitude. Flight condition 7 represents the aircraft during a typical supersonic cruise.

A digital control stability analysis was done for each aircraft axis. A generic z- planerootlocus showing lines of constant damping and lines of constant zeta-omega n appears in Figure9.17. These boundaries willbe used to designate target areas in the z- planein the sections that follow.

9AA Lon_tudinal The unaugmented longitudinal dynamic stability characteristics of theMonarch appear in Table 9.11. As indicated by thesedata,the aircraft has at least one unstable characteristic forevery flight condition. The longitudinal dynamic characteristics required for Level I handlingqualities by MIL-F-8785C (asPer Reference 9.7)appear in Table 9.12. The calculations required for the short period frequency requirements appear in Appendix 3. Figures9.18-9.20 show the open loop root lociin the z-plane, including a target area where the shortPeriodpolesof the system must be placedto achieve MIL-F- 8785C levelI handlingqualities.

Pitchratefeedbackwas used to stabilize theaircraft. Compensating equations were chosen so that theoriginal polesof the open loop system would be cancelled by directly placing a zeroon the calculated polelocation.New poleswere placedin locations in the z-planethatwould give the Monarch level 1 flying qualities. A summary of the compensationequations appears in Table 9.13. The calculations thatdeterminedthese locations appears in Appendix 3. A sampling rateof 100 cycles Per second was assumed from Reference 9.8. The blockdiagram of thepitch SAS appears in Figure9.21. PC MATLAB was used to determinetherootlocusof the discrete system with complete compensation.

Figures9.22-9.24 show theroot loci for thelongitudinal closed loop system for the three flight conditions. An enlargement of theshortperiodpole location has ben included to show itsplacement. As theonly phugoid requirement specifies a damping ratio greater than 0.04,thephugoid rootswere relocated on the stable portion of thereal axisforan equivalent damping ratio of one.

r

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Table 9.11 - Unau_rnented Lon#tudinal Dynamic Stability Charactedstic_ Flight Condition Omega sp (rad/sec) _ _ 1 .12 -.02 -2.42 .81 2 .05 .16 -.36 .14 3 .067 .42 -.40 .15 4 .14 _8 -.413 .172 5 .104 .208 -.588 .289 6 _3 .976 -7.33 .594 7 2.16 .345 -39.63 27.08 8 2.16 .345 -.602 .384 Table 9.12 - Lon_tudinal Dynamic Stability. Requirements S.P. Freq. (rad/sec) _alZLllL,B,i_ Fli__ht Condition Phase Type Min Max Min Max 2 Ground Attack 4.3 16 .35 1.30 4 Combat 3.8 14 .35 1.30 7 Cruise 1.8 12 .30 2.00 0.3 FC 2 Level I Requirements: Zeta = .30 Met _ Not Met 0.2 0.1 x N 0.0 x I (ZetaXorrwga nXT) - / x .g'309 ._66 -0.1 x -0.2 -0.3 • , , I, , I , , , J U,7 0,8 0.9 1,0 l, 1 1.2 REZ Figure 9.18 - Lonotudinal Z-Plane Open Loop Root Locus, Flight Condition ?

0.2 FC 4 Level I Requirements: \.Zeta - 35 Met Not Met 0.1 1%1 o.o ' X (ZeLzXOrnega nXT) = .o_190 .985g x -4).I _.2 , .i , i , i , i V.8 0.9 1.0 l.l 1.2 _Z Fimn'e 9.19 - Lonomdinal Z-Plane Ope. n Loop_ Root Locus. Flight Condition a w FC 7 Level ! Requirement.s: 0.2 Not Met Zeta-30 x Met.

i Z; (Ze_lOmega nXT) = .9630 .gcJ,_l x -0.1 x -0.2 , I = 1 , I , I V.8 0.9 1.0 1.1 1.2 _Z Fimlre 9.20 - Lon_tudinal Z-Plane Open Loop Root Locus. Flight Condition 7 Table 9.13 - Lonotudinal Compensation Equation Summ_,,"y Flight Condition Compensator Numerator (z - 1.1467)(z 2 - 2.0296z + 1.05169)(z 2 - 1.6868z + .717660) (z - 1.1431)(z= - 2.0424z + 1.06452)(z 2 - 1.6868z + .?18413) (z - 1.1467)(z = - 2.0296z + 1.05169)(z= - 1.6868z + .717660) Flight Condition Compensator Denominator (z + .24)(z + .5)(z + .7)(z=- 1.97131z + .973316) (z + .24)(z + J)(z + .7)(z=- 1.97038z + .972003) (z + .24)(z + J)(z + .7)(z=- 1.98839z + .988742)

IRate GZ:o

Fimn'e 9.21 - Pitch SAS Block Diaeram (Zet.IXOn_Kja nXT) = t Zeta - ,30- 0.16 \ 0.14 0.12 0.I0 I%1 ]; 0.08 0.06 0.04 0.02 U.8 RE Z

\

N ]; o Fiexr_ 9.22 - Lon#tudinal Z-Plane Closed Loop_ Root Locus. Flight Conclirlon 2 (ZetaXOrr_o_ nXT] - .9490 .9a59 Ze_- 35 0,14 0,12 FC 4 evel I Requirements Met.

0.I0 K-2.0 0.08 P_a !

0.06 K-0.2 0.04 0.02 K = -0.4 K=O i i i i I i • , 0.00 !

1.00 0.90 0,95 v.85 RE Z N _: o -1 -1 0 1 REZ Fim._ 9.23 - Lon_tudinal Z-Plane Closed Loop_Root Locus. Flight Condidon_ .9944 .9630 (ZelmX_gm nXT] = Zeta 5O I 0.12 I i' 0.10 V V FC 7 Level I Requirements Met v 0.08 I, V K=2.0 V 0.06 0.04 j K=0.2 0.02 / K = -0.I K =-0.4 .0 I.I

z o

K ='0.I Fi_Lrur¢ 9.24 - Lon_tudinal Z-Plane Closed Loop RooI Locus. F_ght Condition 7 The augmented values of short period frequency and damping are shown in Table 9.14. The values of gain were selectedto meet both the handling qualifies requirements and the inertia coupling requirements. Detailson inertia coupling appear in Section 9.5.

Table 9.14 - Au_'nented Lon_mdinal Stability Characteristics Flight Condition .Q_ Omega sp (rad/sec) 2 -0.2 4.35 .38 4 -0.2 3.96 .37 7 -0.1 1.88 .3 Lateral stability in a fighter is very important. If sufficient roll time-constants can not be met, then a rolldamping stability augmentation system (SAS) is necessary. The MIL-F-8785-C requirements, as in Reference 9.7, were examined to determine what the roll and spiraltime constants needed to be. For flight conditions 2 and 4 the Monarch is in flightphase category A, terrain following (TF) and air-to-air combat (CO), respectively. In flight condition 7 the Monarch is in flightphase category B, cruise (CR). The Monarch is considered a Class IV aircraft due to itshigh maneuverability. According to these flight phase categoriesfor a Class IV aircraft MIL-F-8785-C dictatesthe requirements of Table 9.15.

Table 9.15 - Lateral Dynamic Stability. Requirements FC Max Roll Time Constant Min Time to Double Amplitude.

2 1.0 second 12 seconds 4 1.0 second 12 seconds w 7 1.0 second 12 seconds The basic rolldamping SAS block diagram is shown in Figure 9.25. The bank angle to ailerontransfer function was determined using the matrix method of Reference 9.9, using the stability and control derivatives of Section 9.3. The open loop transfer function was determined using the Laplace variables, and then a total pulsed transfer function was determined in the z domain. PC-Matlab was used to perform the z transform.

The z plane root locus was used to find the root locations, and these in turn were used to determine where the spiraland rollroots needed to be. The z plane root locus was then looked at to see whether using a different gain would make a difference. If gain could not solvethe problem then a compensator had to be implemented to move the roots to the desired locations.

T b p ref

[Rate yro

Fibre 9.25 Roll Damping $AS Block Diagram For flight condition 2 it was found that by using a gain of 0.25 that both the roll and spiral time constant requirements for Level 1 could be met. Flight condition 4 inherently met Level 1 handling qualities for the lateral modes, so no stability augmentation was necessary.

For flight condition 7 the uncompensated system was seen to have too small a value of T_2_s. This meant that the amplitude of the bank angle was being doubled too quickly for Level 1 handling qualifies. From the MIL-F-8785-C requirements it was known that the spiral time constant root had to be increased without making the roll time constant greater than 1.4 seconds. The z plane root locus was examined to determine where this was possible. A compensator was then designed to make the sytsem meet the Level 1 requirements. Detail design of the compensator can be seen in Appendix 3. The discrete transfer function of the compensator which is to be implemented in the digital flight control computer is: (z- 1.0009) * (z- .9050) De(z) - (z - .9903)* (z + .I0) The uncompensated z plane rootlocus for flight condition 7 is shown in Figure 9.26. The compensated z planerootlocusisshown in Figure 9.27. A designgainof 0.3 was found to give sufficient roll and spiral time constants for Level 1 handling qualifies.

Detailed development of the roll damping SA$ can be found in Appendix 3. The gains that are necessary for lateral dynamic stability are summarized in Table 9.16.

Table 9.16 - Lateral Control Gains Necessary. for Level 1 Handling Oualities

FC Cain T2s TR

2 0.25 23.1 sec .1327 sec 4 0.0 22.1 sec .4888 sec 7 0.21 28.85 sec .0778 sec 9.4.3 Directional The directional stability of the Monarch will be enhanced with the use of a digital yaw Stability Augmentation System (SAS). The yaw SAS will,when required,improve the dutch roll characteristics of the airplane. Figure 9.28 illustrates the block diagram of the digital yaw SAS system. The unaugmcnted dutch roll characteristics and the corresponding handling level are listedin Table 9.10. The handling level requirements are based on MIL-F-8785C specifications and can be found in Reference 9.7. As stated, only flight conditions2, 4 and 7 are investigated.

For the digital controller, a sampling rate of 100 cycles per second was selected as suggested from Reference 9.8. For fright condition 2 the unaugmented dutch roll discrete root locus is shown in Figure 9.29. The lines of constant danping and constant (oF) for Level 1 requirements are shown. The figure shows that dutch roll Level 1 handling qualities can not be met for any value of gain.

m Detailed development of the compensator for flight condition 2 can be found in Appendix 3. The following implementation equation was developed to achieve Level 1 qualifies for flight condition 2: z= - 1.9978z + .9978 D,(z) = z= - 1.9766z + .9773 m The augmented dutch roll discrete root locus is illustrated in Figure 9.30. For gain ranges of 0 to -1.5,Level 1 handling qualifies arc achieved. A gain of - 0.I is selectedto _== give a dutch roll damping ratio of .60 and a frequency 2.25 rad/sec. These values were m selected to help achieve favorable inertia coupling characteristics as discussed in Section 9.6.

r z, m 0 <K<0.2 0.1 CL._le I "_ = 0.2 Unit ROLL \ i 0,06

.---'-"'--RooTs SPIRAL\

o_, O.OZ

ROOT_ 7

1.02 Z-PLANE REAL AXIS Fi__re 9.26 Uncompensated Roll Z-PlaneRoot Locus forFlightCondition7.

0.1 /--Unit Circle w i 0.06

<_°"

0.02 (.02 Z-PLANE REAL AXIS FiL_ur_ 9.27 Compensated RollZ-Plane Root Locus forFlight Corjdition 7 SERV0 r PLUME DYNAMICS WASHOUT RATE FILTER GYR0 _

I H .... I

Yaw SAS Block Dia_m'am Fi_mar¢ 918 Dutch Roll Characteristics and Corresponding Handling Levels Table 9.17

Ze_D_ *Zt, mt.Qz_a

Flight Condition m,,.

1.5922 .135 2 1 .085 .5665 .104 2 i 2 .183 2.6564 .503 2 3 .1895 4.1945 1.580 1 4 .3769 6.1797 4.040 I 5 .6543 3.1533 .386 I 6 .1223 Two Real Roots outside the unit circle ix, the Z-Plane - Unstable .1299 1.7693 .230 1 As indicatedin Table 9.17, flight condition 4 dutch roll characteristics meet level 1 handling requirements. Therefore no augmentation is required.

The unaugmentcd discretedutch rollroot locus for flight condition 7 is illustrated in Figure 9.31. For increasingnegative gain, it is seen that the roots meet on the real axis and splitto form the oscillatory dutch roll pair. I.,¢vel 1 handling qualities arc still not obtained, as illustrated by the contant damping and (o'I") lines.

Appendix 3 documents the full development of the discretecompensator for this flight condition. The following implementation equation was developed to achieve I.,cvel I handling qualities for flight condition 7: - 2.0088z + 1.0088 D,(z) = z2 - 1.9968z + .99707 The augmented discretedutch roll root locus for flight condition 7 is i].llusu'ated in Figure 9.32. For gains ranging fi'om 0 to 2 itis seen that the dutch rollrolldoes not move much. A gain of 2 is selectedto give a dutch rolldaping ratioof .10 and a frequency of 1.65 rad/sec. These values were selectedto help achieve favorable inertia coupling characteristics as discussed in Section 9.6.

As can be seen from the development of the discrete compensators for the yaw SAS, a different compensator is rexluired for the two flight conditions investigate& Therefore compensator, as well as gain scheduling will bc required. This is possible when using a digital computer to implement the discrete compensator in the flight control system.

t,O II g ' II w w

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S.TX_ .4_u_ otrOlcI-Z 9.5 ROLL PERFORMANCE Roll performance is vital to the success of a fighter. A pilot needs to be able to roll his aircraft rapidly to avoid enemy fire and to point at his enemies so as to lock on ordnances. The maximum roll rate isusually around 150-180 degrees per second, depending on the role of the fighter, according to Reference 9.10. If this roll rate is too high there is the possibility of inertia coupling, and this would keep the plane from being able to roll consecutive loops.

The procedure of Reference 9.7 was used to determine the roll control power derivative due to lateral cockpit control, C_l_del_cpt. This analysis assumes that the Monarch only uses its aileronsfor roll control. The maximum deflection of the ailerons is 25 degrees.

The Level 1 requirements according to M]L_F-8785-C are shown in Table 9.18.

Table 9.18 - Maximum Roll Rate Requirements FC 2: must go through 90 degrees of bank in 1.3 seconds FC 4: for 360 deg rolls: 90 deg in 1 sec, 180 deg in 1.6 sec, 360 deg in 2.8 sec normal flight phase: 90 deg in I.I sex:, 180 deg in 2.2 sec FC 7:50 degrees of bank in 1.1 seconds The equations of Reference 9.7 were solved to find the maximum roll rates of the Monarch. For thc known parameters of each flight condition, the time was put into the equations to sec the level of bank angle response thatresulted. The bank angles thatthe Monarch could go through are shown in Table 9.19, along with the rolltime constants.

Table 9,19 - Bank Angle Response and Roll Time Constants for the Monarch

EC M Ixxs mc.t Tr

2 .65 1067 15370 1 sec 18.7 deg .244 4 .90 678 10631 1.0 sec 282 deg .257 1.6 sec 2789 deg .257 2.8 sec 296213 deg .257 7 1.60 1127 10152 1.1 sec 31.7 deg .264 l The roll performance of the Monarch does not meet Level 1 in all flight phases.

The effect of using the stabilators for roll control for meeting the Level 1 should be investigated. This was not done because it was assumed that the ailerons could provide adequate roll control power. Using the stabilators for roll control would have to be looked at with respect to inertia coupling and roll damping SAS.

w 9.6 INERTIA COUPLING An additional constraint on the selection of gains for theMonarch flight control system was the susceptibility of the aircraft to inertia (roll) coupling. Because of the high roll rates and rapid maneuvers that fighters must execute, these aircra_ are vulnerable to excursions in pitch and yaw while performing combat rolls. Thus, the selection of short period and dutch roll frequencies and dampings were _ated for the three flight conditions to avoid this problem.

The method used for the inertia coupling analysis comes from Reference 9.11.

Plots indicating thevulnerability of theunaugmented aircraft to inertia couplingappearin Figures9.33 - 9.35. Calculation of these data appearin Appendix 3. The width of the "throat" between thetwo hyperbolic boundaries on theseplotsvaries with theproductof short periodand dutch roll damping. The slope of the linewhich starts atthe origin and passesbetween the boundaries isthe ratio of thedutchrollfrequencyto shortperiod frequency.An inertia couplingincident occurs ff thislineintersects one of thehyperbolic boundaries.The roll rateat which thisdeparture occurscan be calculated from the frequencyto rollrate valueson the axes and the corresponding dutchrollor short period frequency. As shown in the calculations in Appendix 3, theMonarch suffers inertia couplingin Flight Condition2 at roll ratesbelow 28 degreesper second. In flight condition 4, theaircraft departs atroll rates below 160 degreesper second. As shown in Section 4.5,therollratecapabilities of the unaugmented aircraft placethe Monarch in the unstable region of the inertia couplingplots.

Figures 9.36-9.38 show the inertia coupling diagrams for the Monarch after implementation of the compensators described in Section 9.4. As seen from the plots, the frequency and damping ratios selected for the Monarch do not produce any instances of inertia coupling. This was made possible by keeping the ratio of short period frequency to dutch roll frequency as close to one as was feasible, pending the restrictions of the handling qualities requirements. The minimum required dutch roll frequency was much less than the minimum required short period frequency for the flight conditions analyzed for the aircraft. As this produced inertia coupling problems, the short period frequency was held at its minimum allowable value and the dutch roll frequency was increased until the ratio of the frequencies moved the line shown in the plots out of the unstable region.

The compensators chosen for the digital flight control system used these frequencies as design points.

rl - 0.85 h- 100 ft.

2 r S - 347.9 sQ. ft.

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_. 0 o 0 2 Ratio of Short Period Frecluency to Steady State Roll Rate (-) Fimm_ 9.38 - InertiaCoupling Bound_-ies for Flight Condition 7 9.7 SPIN DEPAR_ One measure of aircraft spin departure at stall angle of attack is the value of Cn beta dynamic. Using the methods of Reference 9.11, the aircraft has adequate resistance to dcparan-c when the sign of Cn_beta dynamic is positive. Calculations of this parameter appear in Appendix 3. The results arc summarized in Table 9.20. These data indicate that the Monarch does not enter a divergence at the stall angle of attack. This analysis gives no indication of the spin tendencies of the aircraft in the post-stall flight regime.

Table 9.20 - Results of Spin D_parture Analysis Flight Condition _ M Cn beta dynamic 1 0 .20 .4532 2 100 .85 .0955 3 10000 .90 .2099 4 15000 .90 .3789 5 30000 .90 .4770 6 30000 1.20 .3285 7 30000 1.60 .0908 8 40000 .80 .2291 According to Reference 9.11, a means of aiding aircraft spin recovery is to locate the most aft center of gravity of the aircraft ahead of the centroidof the total aircraft planform. This guarantees a form of stability margin at an aircraft angle of attackof 90 degrees. As shown in Figure 9.39, the most aft center of gravity is ahead of the planform centroid. This margin of stability for the aircraft is 8.12 inches,or 5.64% of the mean geometric chord. Therefore, these analyses indicatethat the Monarch does not have any inherentspin tendencies.

9.8 LOW LEVEL RIDE O_UALITIES Attack mission require fighter aircraft to fly low level, high speed profiles to the target. The aircraft and the pilot must be capable of accurately delivering ordnance in this flight regime. An assessment of the low levd ride qualities of the Monarch was completed to determine ff the aircraft reXlulrrA a ride quality augmentation system.

The method for this analysis comes from Reference 9.7. A "root mean gluare, d g- level"Per foot per second gust level ('_) was calculated for the aircraft in Appendix 3.

.=.

While thisanalysisis usually done only for low level,high-speod flight, the values of A .:- ,.

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Table 9.21 - Monarch Ride Qualities Flight Condition h. ft. M _ 1 0 0.2 30744 .00497 2 I00 0.85 28395 .027774 3 I0000 0.90 23980 .02742 4 15000 0.90 23980 .02385 5 30000 0.90 28223 .01546 6 30000 1.20 24534 .01818 7 30000 1.60 27239 .01465 8 40000 0.80 29917 .009324 u Values of A greater than 0.005 generally indicate deficiencies in ride qualities. This analysis indicates that the aircraft may need a ride quality augmentation system throughout most of its flight envelope. The extensive augmentation of the standard flight control system may provide some improvement in the ride qualities of the aircraft, although this would have to be verified in flight test.

9.9 VERTICAL TAIL/RUDDER REMOVAL STUDY A study was done on the Monarch to replace or reduce the vertical tail of the airplane with a thrust vectoring control that would provide partial or full directional stability. Eliminating the vertical tail would reduce the radar signature of the aircraft, reduce or eliminate interference drag between control surface at the aft end of the aircraft, and may reduce the complexity of some of the flight control system.

The preliminary design of the Monarch used a vertical tail with a 63 sq. ft.area.

Using the lateral gust handling qualities requirements of MIL-F-8785C, which allow for a w 30 knot gust, equivalent sideslip angles were produced for the eight flightconditions developed for the stability and control analysis. A proposed thrustvectoring location(FS 625) was used to determine: * the sideforce that would be needed to control an aircraft without a vertical tail, * the sideforce that would be needed to control an aircraft with the original vertical tail but without a rudder.

The calculations for this analysisappear in Appendix 3. The results am summarized in Tables 9.22 and 9.23.

Table 9.22 - $ideforce Required for Adequate Stati_ Directional Stability.

Flight Condition _ Sidcfor_e at FS 625. Ibs 1 0.2233 2583 2 0.0534 13077 3 0.0522 9712 4 0.0532 8038 5 0.0566 4231 6 0.0424 5926 7 0.0318 7939 8 0.0654 2339 Table 9.23 - Sideforces Needed to Replace the Rudder Eligh.L.Qo/_ $ideslip. rad Rudder Defl..deg $ideforce _ FS 625. lb 1 0.2233 -0.478 -61.5 2 0.0534 0.309 962.8 3 0.0522 1.49 4318.3 4 0.0532 1_8 4219.5 5 0.0566 2.19 2663_ 6 0.0424 1.83 4232.7 7 0.0318 -0.161 ..441A 8 0.0654 1.47 729_ In the case of vertical tail removal, high dynamic pressures in flight condition 2 produced unreasonably large sidcforces to compensate for the lack of inherentdirectional stability. The design for the yaw thrust vanes for the Monarch, as shown in Chapter 6, Figure 6.11, did provide adequate directional stability for removal of the rudder. The I correlation between sideforce and yaw vane deflection appears in Table 9.24. The physical limit of the yaw vane deflection was 25 degrees. The calculations appear in Appendix 3.

Table 9.24 - Sideforce8 Produced by Yaw Vane Deflections Flight Condition Sideforee at FS 625. lb Vane Deflection Angle. deg.

1 61.5 1 2 962.8 4 m 3 4318.3 12 u 4 4219.5 6 5 2663.2 I0 m 6 4232.7 17 B 7 441.1 1 8 729.9 9 N Additionally, yaw vane deflections a.Uowed for a reduction in the original vertical tail area of the Monarch. The tail area was reduced from 63 square feet to 40 squar_ f_t.

r_ B The calculation for this analysis appear in Appendix 3. This empennage configuration was successfully integrated into the dircetional stability analysis of Section 9.4.

i REFERENCES FOR CHAPTER 9 9.1 Roskam, Jan, Ah'p. lane DesiL_. Part VI: Preliminary. Calculation of Aerodynamic. Thrust and Power Characteristics, Roskam Aviation and Engineering Corporation, Route 4, Box 274, Ottawa, KS, 66067, 1987.

9.2 Cox, Brian, et.al., Preliminary. Stability and Control for the LIFT Confimn'ation.

Universityof Kansas, AE 622, 1990.

9.3 Roskam, Jan, Airplane FlightDynamics and Automated Controls. Part H, Roskam u Aviation and Engineering Corporation,Route 4, Box 274, Ottawa, KS, 66067, 1979.

9.4 Hoak, D.E., et al, USAF Stability. and Control Datcom, Flight Control Division, Airforce Flight Dynamics Laboratory, WPAFB, Ohio.

9.5 Pratt and Whitney Advanced Engines Performance, Weights and Dimensions Model, Fighter/Attack/Interceptor Application, Provided by Public Release from Wright Research and Development Center, January 1990.

i 9.6 Northrop Aircraft, Inc., "Dynamics of the Airframe Addendum", September 1952.

9.7 Roskam, Jan, Airplane Desi_. Part VII: Determination of Stability.

Control and Performance Characteristics: FAR and Military. Requirements, Roskam Aviation and Engineering Corporation, Route 4, Box 274, Ottawa, KS, 66067, 1988.

9.8 Personal Conversation with Dr. Jan Roskam, April 24,1990.

9.9 Roskam, Jan, Ah"p.lane Flight Dynamics and Automated Controls. Part I, Roskam Aviation and Engineering Corporation, Route 4, Box 274, Ottawa, KS, 66067, 1979.

9.10 Whitford, Ray, Design for Air Combat, Jane's Publishing Company, 238 City Road, ta,J London EC1V 2PU, 1987.

9.11 Roskam, Jan, Airplane Design. Part m: Layout Desi_ of Cockpit; Fuselage.Wing and Em_nnage: Cutaways and Inboard Prof_ Roskam Aviation and Engineering Corporation,Route 4, Box 274, Ottawa, KS, 66067, 1986.

D 10. STRUCTURAL LAYOUT, MATERIALS SELECTION. MANUFACTURING PROCESS, ACCESSIBILITY AND MAINTAINABILITY CONSIDERATIONS The purpose of this chapter is to present the preliminary smacmral design and L m materials selection for the Monarch aircraft. The manufacturing process and accessibility and maintenance considerations are also presented here. Structural design is based on the Class I methods presented in Reference 10.8 and a data base of current fighters. Materials selection is based on a desire to create opportunities for weight and life-cycle cost reductions. The manufacturing process is presented with a shop flow diagram and a description of the processes used in forming the aircraft. The accessibility discussion presents the engine removal schemes and other systems maintenance considerations.

10. I. STRUC'TURAL ARRANGEMENT The purpose of this section is to present the preliminary su'uctural arrangement for the Monarch aircraft. The work presented here is used to indicate where primary structural w members are located to provide stiffness and component mountings for the Monarch. More advanced structural design requires information on aircraft loads. Loads information was not calculated for the Monarch aircraft. For this reason the structural arrangement of the Monarch is currently based on data for structure of other fighters and assumptions of primary load paths. The design considerations and the comparative data base for each primary structural component are also presented.

10.1.1. Fuselage Structural Arrangement The driving design considerations for the fuselage structure were to reduce the number of primary frames and achieve synergism wherever possible. Materials used for frames and longcrons are aluminum and titanium. The skins use aluminum and composites. The location of secondary frames and longerons was based on the fuselage w layouts in Reference 10.12.

The location of major frames for the Monarch is shown in Figures 10.1 and 10.2.

These frames are made of titanium. Lesser flames are made of aluminum. These flames are spaced at intervals of 18 inches aft of FS 116.5. Synergism was achieved at the major fuselage frames as shown in Table 10.1. Note the location of jack points on frames FS 364 and FS 552. These points are used during the production stage for the testing of systems and can be used during service life as securing points during tire and landing gear replacement.

u Longerons are also used to stiffen the fuselage and support components. Major longerons are placed so that landing gear bays, the weapons bay, canopy, and nozzle openings receive large amounts of local stiffening. Lesser longerons are placed along the aircraft at 12 inch intervals.

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o m w (S3H_)NI) NOll¥.LS 3NI'IEI3.LYM T_c 10.1 FLL_Cla_e Faro'he Syner_sm 9. FS 412 1. FS 116.5 Radar Mount Aft Weapons Bay Inlet Support Nose Gear Support Fuel Cell Support Canopy Attachment I0. FS 421 2. FS 131 Wing Attachment Point Support Nose Gear Bay Forward Pressure Bulkhead Inlet support Fuel Support II. FS 437 3. FS 195 Nose Gear Attachment Pt Wing attachment Pt.

Seat Attachment Cockpit Support 12. FS 461 4. FS 212 Inlet/Compressor Interface Front Engine Mounts Front Main Gear Bay Support DL Inlet Support Main Gear Attachment Pilot Armor/Insulation Rear Pressure Bulkhead 13. FS 493 5. FS 250 Shelf Attachment Aft DL Engine Mounts Front Weapons Bay 14. FS 533 6. FS 288 Main Engine Mounts Inlet Support Shelf Support Fuel Bay support Gun ring Supports 15. FS 552 7. FS 297 Vertical TailMount AIM-7 Mount Rear Main Gear Bay Support Inlet Support fuel Bay Support 16. FS 578 8. FS 350 Vertical Tail Mount Aft Gun Mount Ventral Nozzle Mount Aft AIM-7 Mount Shelf Support Fuel Support Wing Shelf Support 17. FS 612 Vertical Tail Attachment Horizontal Tail Attachment Engine Slip Mount 10.1.2. Wing Smactural Arran eement The design drivers for the wing structure of the Monarch are: * The ability to sustain air loads * Fuel storage volume * Lack of wing carry through * Weapons station requirements * Control surface requirements * Weight considerations The wing structure must be able to support sustained loads of 9g's. The wing is also subjected to fatigue due to gusts and loads caused by deflections of the high lift devices. Locations for weapons carriage must also be provided.

The wing of the Monarch is to be used for fuel storage. This requires that the volume of the structure be kept to a minimum. Additional structure such as baffles and allowances for fuel tank access must also be made. These requirements act contrary to the requirement for a minimized component weight.

Wing spar carry-through was not possible as the Monarch is a mid-wing configuration and carry through spars would conflict with the engine section. This required that additional support be provided where the wing joins the fuselage. Lack of wing carry- though is not uncommon in fighters (see Table 10.2), but it does result in a weight increase at the fuselage/wing interface.

Table 10.2 presents data for wing _s used in other fighter aircraft. This information "was used to dem'mine _hc strucucal layout for the wing of the Monarch.

Actual sizing of the wing members is not possible until loads are calculauxt.

The wing structural layout of the Monarch fighter is shown in Figure 10.3. The structure consists of four spars, seven ribs, and a "shelf'. The number of spars is less than that used by most fighters. This is assumed possible through the use of titanium spars and highly stressed skins. The wing attaches to a "shelf' much in the manner of the F-16 (see Figure 10.4). Structural components indicated with letters in Figure 10.3 collectively make up the shelf of the Monarch. Spar attachment points and rib locations are given in Table 10.3. Ribs are used as divisions in the fuel tanks as well as to provided stiffness.

Weapons hard points are installed at B.L. station +/-87 and +/-135 at fuselage stations 470 and 490, respectively (Reference 10.9).

Structural synergism was achieved at rib numbers 3, 5, and 7, and at spar numbers 1 and 3. Rib numbers 3 and 5 act as weapons hard-points as well as wing stiffeners. Rib number 7 act as both the spar cap and the mount of the AIM-9 launch rail. Spars 1 and 3 support the wing and provide mounting locations for control surfaces.

The materials used in the wing arc titanium and composites. The upper wing skins u arc made of graphite epoxy. The spars and ribs arc made from Ti-6A1-4V titanium alloy.

The lower skins are made from boron epoxy. The leading edge devices are made with aluminum skins and an aluminum honeycomb core. Trailing edge devices are graphite W < _J W bD D tO u.

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Justification for the selection is provided in section 10.2.

Table 10.2 Fighter Wing Structure Data Aircraft Spars Ribs Carry Through F-15 5 I0 No F-16 9 11 No AV-SB 8 6 Yes MiG-21 m 5 NA No MiG-23 4 12 No MiG-25 4 4 No A-4 3 7 Yes F/A-18 6 NA No w Source: References 10.2 and 10.8 z w w = Table 10.3 Wing Structural Layout Spar Shelf Attachment WI, Station

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1 425 160 2 450 160 i 3 475 160 4 500 160 Rib # BL Station WL Station Shelf +/-44 160 1 +/-61 160 2 +/-77 160 3 +/-87 160 4 +/-107 160 5 +/-135 160 i 6 +/-171 160 7 +/-202 160 10.1.3. Horizontal Tail Structure Design drivers for the horizontal tail were the desire to reduce weight, increase heat resistance, and create structural synergism. Weight reduction is achieved through the use of composites and a sandwich/core structure. Synergism was achieved by placing the horizontal taft attachment points at the same fuselage fi'am¢ as the vertical tail and engine slip ring.

Table 10.4 presents a data base of aircraft which use differential stabilizers. Based on this information, the structural layout of the horizontal tail was chosen.

Table 10.4 Fighter Stabilizer Dam Aircraft Spars Ribs Material F-111 5 4 Boron Epoxy Mig-23 3 8 NA F-14 NA NA Boron Epoxy F-15 2 3 A_ core Graphite Epoxy Skins NA F/A- 18 NA A1 core Graphite Epoxy Skins Source: References 10.1, 10.2, and 10.3 The horizontaltail structural layout of the Monarch fighter is shown in Figure 10.5.

The structureconsistsof two spars and three ribs. The tail attaches to the fuselage at the second ,spar as indicated Table t-0.5. The fast spar acts as a re-enforcement for the leading edge and an attachment point for the fixed slat. Rib number one acts as a local stiffener at the attachment point. Rib number three acts as a spar cap. The basis for this layout is the 1::-15 which uses an almost identical arrangement (Reference 10.2).

The I=-15 uses an aluminum honeycomb core with graphite epoxy skins (Reference 10.1). Aluminum spars and ribs, and a core of titanium honeycomb are proposed for the horizontaltailof the Monarch. The skins are made of a carbon/carbon composite. A justification of the materials selection is presented in section 10.2. A diagram showing the actuator mechanism for the horizontaltailis shown in Figure 10.6 and disucssed further in Chapter 11.

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Table 10.5 Horizontal Tail Structural Layout WL FS BL Root chord Spar # Station Station Station Location (inches) (inches) (inches) (%C, inches) 1 160 582.7 45 10% (6) 2 160 606.7 45 50% (30) Rib # WL FS BL Span Station Station Station Location o_gm 1 160 587 45

7.5%(5)

2 160 595 55

9% (6)

3 160 648 I15 100% (65) 10.1.4. Vertical Tail Sta'ucnmd Arralagement The purpose of this section is to present the vertical tail structmal arrangement for the Monarch aircraft. The design considerations and a data base for comparison are presented. The structural layout, materials used, and structural synergism are also given.

The primary design drivers for the vertical taft are: * Ability to sustain air loads and provide control.

* Battle damage tolerance * Structural Synergism * Ability to store antennas, IFF gear, etc.

The vertical tail must be able to withstand sustained and repeated combat air loads.

Fatigue due to gusts is part of this consideration. The vertical tail structure should also be able to tolerate the loss of some members without catastrophic failm'_. This requires that the structure be formed to transmit loads around severed members, or that members be made redundant. The second method results in a weight penalty. The fast method requires additional det_!ed design. It was assumed that other fighters must meet these same requirements, so that by using a similar structure this requirement would be met for the Monarch. This design was coupled with composite skins which transmit loads around damage to create better short term damage tolerance than metal skins (Reference 10.6).

Synergismfor thevertical tail was achievedby placing itsspars such thatthe ribs support: * Chaff and flare dispenser * Fuselage frame slip ring * The horizontal tail attachtmnt The size of thevertical tail results in useful internal volume. Synergism is added by usingtheroom to store components such as IFF and radioantennae in the vertical tail.

This allows thereceivers to be located away from theinterference caused by aircraft systems.

Table 10.6 presents a database of the vertical tail structures used in otherfighter aircraft. The aerodynamic loads that the vertical taft will experience have not yet been calculated. For this reason, the vertical tail smacmre is based on this data.

The vertical tail szrucmral Iayout of the Monarch fighter is shown in Figure 10.7.

The smacture consists of four spars and four ribs. This selection agrees with the data base.

The tail attaches to the aircraft at the location indicated in Table 10.7. Attachment is to the "spine" longeron and fuselage frames. The ribs provide for tail stiffness at mounting points, an equipment mounting shelf, and as a means for loads to be nmasmitted in case of spar failure. Rib number four acts as a way to dissipate lightning strikes. The spars provide stiffness and redundancy for battle damage.

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_u o Lq w o 0 0 0 o U1 0 I_ m r_l e_J (53H3ND NOI.I.Y.LS 3NI'I_3.LVM Table 10.6 Fighter Vertical .Tail Structure Data Aircraft Spars Ribs Material F-4 3 NA Ti,A1 F- 15 2 1 Ti, Boron Epoxy F- 16 4 3 AI, Graphite Epoxy F/A-18 6 0 Ti, Graphite. Epoxy _-tt t ,6 7 A1, Gralflfite Epoxy Source: References10.1 and 10.2 The primary materials used in the vertical tail arc 2024 Aluminum, Ti-6A1-4V titanium alloy, and graphite epoxy. Aluminum is used from the leading edge of the vertical rail to spar #i. Titanium alloyisused in the attachment points, spars, and ribs.

Graphite epoxy skin isused with the exception of theleadingedge.

Table 10.7 Vertical Tail Structural Layout Spar # WL Station FS Station Root Chord Location (inches) (inches) (%C, inches) w 1 192-291.8 563.5 15% (16.2) 2 192-291.8 581.8 32% (34.2) 3 192-291.8 600.1 48% (52.2) 4 192-291.8 618.4 65% (70.2) Rib # WL Station Span Location Function (inches) (%b, inches) I 191 0 (0) Attachmentpoint 2 221 30% (30) Equipment Shelf 3 234.3 66% (66) Stiffness 4 291.8 100% (100) Lightning Dispersal 10.2.MATERIALS $ELRCTION. JUSTIFICATION. AND LAYOUT The purpose of this section is topresent the materials layoutand selection justification forthe Monarch. Design criteria for materials selection are given by Reference10.7 as: * MechanicalPrope_es -Static Su-ength Efficiency -Fatigue -Fracture Toughness and Crack Growth -Environmental Stability * Fabrication Characteristics -Availability and Productibility -Material Costs -Fabrication Characteristics Other considerations for theMonarch areweight savings, damage tolerance, and cost.

Weight savings are achieved throughthe use of compositematerials and materials with high strength to weightratios. Damage tolerance is achieved by usingmaterials that have high toughness and redundant structure. Damage tolerance is further increased by using structural methods that redistribute loads well. An example of thismethod is sandwich/honeycomb structure which isused in the nailingand leading edge surfaces.

Many of the materials selected for the Monarch have high initial costs. However, it is possible that these costs are regained through the life cycle of the aircraft by better performance. References 10.4, 10.5, and 10.14 indicate that materials such as composites and titanium offer better fatigue characteristics, weight reductions or both. The materials of the Monarch are chosen to be light and have good fatigue properties. If it is possible to capitalize on these properties, life cycle cost may be lowered through reduced fuel consumption and maintenance requirements. The materials of the Monarch have been selected with thesepossibilities in mind.

This section isdividedby materials type. The location of mamrials is as indicated in Figure10.8. Table 10.8and Figures 10.9-12provides a database forcomparison of materials usage in the Monarch and current fighters. The materials used in theMonarch w are as follows: Aluminum Alloys This alloy is used primarily in thefuselageand horizontal tail.

Fuselage Frames: Those frames which are not in engine heat fields or heavily stressed are made of 2219 A1-Cu. 2219 A1-Cu is relatively tough and resists corrosion cracking well (Reference 10.4). It has a yield strength of about 60,000 psi (in tension), and resists creep. 2219 does not retain strength well above 200 F. Manufacun'ing is relatively easy and a large number of suppliers exist.

Table 10.8 Aircraft Materials Breakdown Aircraft AI Steel Composites Ti Other HIMAT 25 9 29 19 18 B-52 69.7 11.5 1.6 1.5 15.7 F-14 36 15 4 25 20 F-15 37.3 5.5 NA 25.8 NA F-16 80 8 3 1.5 7.5 F/A-18 49.6 12.9 9.9 12.9 14.7 AV/SB 47.7 NA 26.3 NA NA All values in percentage of aircraft take-off weight Note: Data for AV/SB is for skins only Source Reference 10.1 and 10.2 ;,,=; w w w .= n i limi ORIGINAL PAGE IS OF POOR QUALITY w

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¢q Horizontal Taft: 2219 Al-Cu was also used in the spars and ribs of the horizontal taft. This material was chosen due to its resistance to creep and ability to tolerate temperatures up to 600 F (Reference 10.4). Although exhaust air through the yaw port is in the vicinity of 1400 F, loss of strength due to the heat effects of the yaw ports was not considered detrimental. This assumption was made due to the fact that the yaw ports are open only forshort Periodsof time (see Chapter7),the coolingeffect of freestreamair, and protection by the skins.

Stringers: 2024 was used in fuselage stringers. This material was chosen on the basisof its ability to be heat u'eated to high strengths and toughness, its cost, and availability. 2024 isalso creep resistant at elevated temperatures (Reference10.7).

FuselageSkins: Large portions of thefuselage skinsneed to be aluminum due to heatfields and the possibility of foreign objectdamage. The exactalloy willdepend on thelocalheat fields and strength requirements.

Titanium Alloys: Titanium is used in the wing,vertical tail, engine section, horizontal tail, and fuselage frames.

Wing: Titanium is used in the wing spars and ribs. This material was chosen based on its high strength-to-weight ratio (1.3 that of AI, Reference 10.4) and good corrosion and fatigue characteristics. Titanium also has a low thermal expansion coefficient. This allows good bonding atmetal/composite interfaces. Wing/Fuselageattachment points are made of Ti-6AI-4Vpreparedusingpowder metallurgy techniques.Reference10.5indicates w that exceptional fatigue and crack stoppageispossible with thismaterial.The penaltyis a very slight decreasein tensile strength.

w Vertical Tail: Ti-6AI-4V isused for thesparsand ribs for the same masons given in thewing description.

Engine Sections: Ti-8AI-Mo-IV is used in the structure surrounding the engine section of the Monarch. This material offers good creep and thermal stability characteristics up to 850 F (Reference 10.5). An additional benefit is that the material can be welded.

Horizontal Tail: A titanium honeycomb is used as filler between skins. This application offers high strength and good heatresistance. This style of application ismore tolerant of battle damage than that withonly spar/skin arrangements (Reference10.6).

Fuselage Frames: Ti-6A1-4V is used in fuselage frames for attachment points. This choice was based on the strength-to-weight ratio of this material.

Steels: Steelalloys areused in various applications including landinggear, fasteners, and other components which require high strength.

Composites: Various composite materials are used in the skins of the Monarch fighter. In general, composites arc light, strong, and have good corrosion resistance. Methods have been developed by McDonnell/Douglas for repairingbattledamage to composites and the AV/SB uses large amounts of graphite/epoxy. A precaution against lighming and bird strikes exists in the form of metal leading edges. Given that composites are able to operate with field repairsand tolerate the heat fieldsfound in hover, the precautions mentioned should make composite materialsextremely serviceable for the Monarch.

Vertical Tail Skin: Graphite Epoxy is to be used for this applicationto achieve weight savings and high strength.

Horizontal Tail Skin: A carbon/carbon composite skin is used to provide heat resistance at low weight. Heat resistance is required due to the location of yaw RCS ports.

Wing: Boron Epoxy composites are used on the lower surface of the wing due to high strength/weight values in tension,heat resistance, and corrosion resistance.Graphite Epoxy is used in the upper wing surfacesbecause it is cheaper and heat resistance is not so crucial. Note that Boron poses some environmental problems which have been considered.

Fuselage: Graphite Epoxy is used in all access panels as a method of weight savings.

Other Materials: Various other materialsare to be used in the constructionof the Monarch. A non- exhaustive list of examples includes: w * Rubber (tires) * Plexiglass(canopy) = * Composites (radome) * Rene 41 (nozzles) 10.3. MANUFACTURING BREAKDOWN AND PROCESS The purpose of thissection is to present the manufacturing process and shop flow for the Monarch fighter. The manufacturing breakdown is shown in Figure 10.13. The shop area required for the Monarch is presumed to be 50,000 sq. ft. based on comments fzom Reference 10.15. This area is to house all stages of production at a peak production rate of 10 aircraft per month.

Several different processes will be used in the manufacturing of the Monarch. Raw materialswill be received in a storage and testing area where a qualitycontrol group will verify that the materials meet specifications. Required materials are then requisitioned from this stockpile as needed. Fuselage frames are milled using computer controlled milling machines, or forged and heat treated. Wing, vertical taft, and horizontaltaft su'uctures are created from standardbar stock. This stock is formed using a number of methods includingmilling, rolling, and drawing. It may be possibleto purchase the wing and tail L su'ucturcs in finished, unassembled form from subconwactors. Fuselage skins are cut from _ FOLDOUT FRAME / CA_O_LAUG£ A_IO WEAPONS t WING ASSEMBLy ENOiNES NOZZLES L FOLDOUT FRAME RADOME CANOPY WING STRUCTURE _ NOSE ASSEMBLY WING SKINS FGRWARO FUSELAG£ ONTROL SURFACE5 <c MID-FUSELAGE LANDING GEAR V TAIL STRUCTURE H TAIL ASSEMBLY H TAIL STRUCTURE ,na.rch Shop Row _ <2'_ -, H TAIL SKINS sheet stock and formed using stamping or plastic forming methods. Wing and taft skins are made from autoclaving methods described in Reference 10.16. Due to the precise requirements and conical nature of the radome, it is created using a computer controlled filament winding process. Metallic skins will be joined to the fuselage using rivets.

Composites will be joined to the structure using titanium/composite lap joints or by riveting through re-enforced holes. Other components such as the canopy, landing gear, and engine are supplied by subcontractors.

The shop flow of the Monarch follows a path consisting of: I. Production of structure for allcomponents.

2. Installation of systems in aft and mid fuselage sections.

3. Assembly of forward fuselagesectionincluding ejection seat, canopy, and skins.

4. Joining of fuselagecomponents.

5. Installation of landing gear and systems.

6. Applicationof fuselage skins.

m.a 7. Addition of vertical and horizontal tailassemblies.

8. Addition of wing assemblies.

9. Installation of engines and nozzles.

w 10. Final systems check.

11. Applicationof camouflage paint and addition of armaments.

At each step in the manufacturing process qualitycontrol measures should be taken toinsurc that all work has been performed correctly. Two types of quality controlare currently being used in industry. The most dominant type of quality controlestablishes a separate deparmaent outside of manufacturing to perform checks. This tends to create resentment among the people responsible for manufacturing and adds to clerical and accounting cost by requiring a separate department and staff. An alternative method of quality control which is beginning to appear in U.S. industry is "Total Quality Management" or TQM. Under TQM, the manufacturing group is directly responsible for the quality control process. This reduces staffing costs, reduces worker tensions, and may instill better workmanship by making each individual directly responsible for his work.

TQM does suffer the drawback of removing objectivity in the person checking the work.

One compromise between these two systems may be to integrate quality control specialists directly into manufacturing groups. These specialists would be able to retain their objectivity as they would be checking the work of others, but they would also be an integral part of the team rather than an outsider. This would reduce clerical costs and departmental conflicts.

10.4. ACCESSIBII.rI'Y AND MAINTAINABILITY CONSIDERATIONS The purpose of thissection is to present the design considerationsaffected by accessibility and maintainability. These two factors heavily influencethe lifetime costs and combat success of a fighter.Easy maintenance reduces manpower costs and increases combat effectiveness by allowing quick repairof battle damage. Examples at the extremes of thisscale are the F/A-18 and AV-8B. The U.S. Navy record for engine removal and replacement on the F/A-18 is eight minutes under competition conditions. In comparison, the entirewing of the AV-8B must be removed in order for the engine to be changed.

The accessibility and maintainability features am divid_ between engine removal and systemconsiderations. Only major accessportsare mentionedhem as examples.

Mentioning all accessrequirements isnot particularly usefulat thisstageof the design and would be extremely complicated.For example, some 80% of the skins on the F-15 am accesspanels.

The fonowing discussion focuses on emergency accesstoprimary systems. During

the life of theMonarch, several complete overhauls will have to be performed under depot conditions. The work performed duringtheseoverhaulswillbe sp_ificdby military regulations and willchange as the aircraft ages. For thesemasons, depot maintenance is not addressed in this report.

10.4.1. En_ne Removal r The engine removal ispresented by showing a stepby stepprocedure. Engine r removal considerations arc: w * Engine accessibility from ground level * Sn'uctural soundness * Accessibility in all typesof NBC and Arcticgear Removal of the lift engineisslraight down through the nozzleopening. This avenue was chosen totakeadvantage of an existing structural opening,and to avoid breakingframes or disconnecting non-engine systems. For similar masons, the cruiseengine isremoved in the aft direction. An engine removaljack isshown in Figure 10.14. This jack is a preliminary designdrivenby the following considerations: * Able toremove both engines * Able to _ over roeg-h ground * Self-powerezl * Remotely conm_llcd The ability to remove both engines is provide, d in thefork-lift type arrangement for the lift engine and theupper rails for the cruise engine. Rough ground operations arc achieved througha wide wheel base and large, soft tires. The jack is to be powered by a 300 hp.

diesel or gasoline engine. This size isestimatexl to be adequatefor the powering of all jack systemsand ground _'ansportation of both engines.Remote control is desirable as it allows themechanic to _osition thejack in the correct relationship with the removal deviceswithoutrequiring thatdirections be relayed through a second party. This is accomplished by connectinga hand helddeviceto the engine conu'ols throughelectric cords.

Figures10.15 and 10.16 show the mountings and accesspanelsfor the lift engine.

Removal procedure forthisengine is: I. Disconnect enginefuel and coolantsystemsthroughaccesspanel "I."

2. Remove access panel "2" around the enginenozzle.

FOI,,,I_OUT FRAME: /.

NOTES: ALL DIMENSIONS IN INCHES POWER SUPPLIED BY 300 HP AUTOMOTIVE ENGINE ALL MEMBERS CURRENTLY 1.5 IN STEEL BAR 29.3

f

w "'--4 SCALE: 1150 = 167.5 ,,- r 36.0 • ..,,._ C_.C_ r' I ,_.° .

f Lo.o

Fimn'e 10.14 En_ffn¢ Re

FOLDOUT FRAMEc_-

REPRESENTATION OF ENGINE AND DRIVE TRAIN

302.5 =4

, 250.0 :--i "m z _w ,r_t: =tz i- .< w<< <zo m L m

/\

\

2O0 3. Use the engine removal jack to take the weight of the lift engine off engine mounts EM1-4.

4. Undo the inlet/compressor interface latches through access panel "3P" and "3S."

5. Remove bolts EB1-4 from engine mounts EM1.-4 through access panels "1", "2", "3P", and "3S."

6. Lower engine and nozzle through access panel "2" using the engine jack.

This procedu_ is shown in 10.18. The engine is replaced by following the steps in reverse procedure. Note that access panels and bolts are sized so that they can be removed in arctic or other protectiveclothing. Access panels can be reached without the aid of the lackiers (the highest reach required for eitherengine is 6.25 ft.).

Figure 10.15 shows the engine mountings for the cruise engine. Removal follows as: 1. Disconnect engine systems and pitch RCS ports through access panels through main landing gear bays (see Figure 10.17).

2. Disconnect fi_"A/compressor interface latches fllrough landing gear bays.

3. Disconnect ventralnozzles through access panels "4P" and "4S."

4. Disconnect RCS yaw ports through access panels "5P" and "5S."

5. Remove access panel "6" around cruise engine nozzle and disconnect the nozzle.

6. Align the cruise engine removal jack rails with the engine rafts of the aircraft.

7. Activate the grappling system of the removal jack so that itgrips the removal posts at the nozzle/engine interface.

8. Disconnect the main mounts through landing gear bays.

9. Disconnect the engine slip ring through access panels "5S" and "5P."

I0. Reverse the grappling system so that the engine is drawn out of the aircraft, along the engine rails, and onto the removal jack rail This process is diagrammed in Figure 10.18. Again note that the panels arc sized so that they can bc used in all types of protectiveclothing. The highest point a mechanic must be able to reach is 6.25 ft.from ground level.

E

© _

i

L_ u_ ,IC

©

LL_I

_Z

FOLDOUT FRAME /.

--ACCESS PANELS u

/7

14_I

I , I l,ol ,ol

l,e

©

i STEP I DISCONNECT SYSTEMS: FUEL w CONTROL NOZZLE FAIRING INLETS !

w i lli- mi_ STEP 3 LOWER ENGII REPLACE EN Fi_,m_ 10.18 Lift Engine iiii, FOLDOUT FRAME c>_ NOTE_ APPERANT CONFUCT BETWEEN FUSELAOE AND CART ODES NOT OCCUR AS CAN B_. SEEN FROM A TOP VIEW

-

STEP 2 USE ENGINE JACK TO TAKE ENGINE WEIGHT REMOVE ENGINE RE"I'A[NING BOLTS THROUGH ACCES PORTS E AND REVERSE AWAY INE USING REVERSE METHOD

Er, m.oz

FOLDOUT FRAME / ACCESS PANELS i

© ©

STEP ! DISCONNECT SYSTEMS: FUEL CONTROL NOZZLE FAIRING INLETS

o ©

STEP 3 DRAW ENGINE AFT 0 REVERSE AWAY FROI -k , = = 10.19 C w FOLDOUT FRAME ¢:_

/7

© 0

STEP 2 ALIGN ENGINE JACK WITH REMOVAL RAILS ENGAGE GRAPPLING MECHANISM

/7

. I _F-i i i 111 !F-I, TO CART AND AIRCRAFT ise En__e Removal 2O4 10.4.2 Systems Access Primary systems access panels are shown in Figures 10.16 and 10.17. The numbers in these figures correspond to indicated panels in Table 10.9. Primary systems are access_ through the following panels as indicated in Table 10.9. Note that some systems access recluires removal through the top of the aircraf_ Based on systems placement, this was unavoidable, but it is not an i_al practice. Such access requires addi6onal equipment such as ladders which increase cost, complcxiw, and maintenance time.

w 2O5 Table 10.9 Systems Access Panels

Frame #

Access through side 7 1 Radar Port side, gun drops down 8 15, 19 Gun Starboard side 9 15, 19 Ammunition Remove on trays 10 21, 28 Avionics Accessed from above 11 40, 41 ECM Lower 12 Wing spar 1 LE flight wing surface controls Lower 13 Wing spar 2 'rE flight controls wing surface Through landing 21, 23 Flightcontrol gear bay motor Through landing 26, 28 Flight control motor gear bay Remove downwards 14P,S 35, 37 Stabilator Actuators Remove upwards 15 Hydraulics Remove aft, down 16 32, 40 APU Remove upwards I7P,S 33, 35 Electric Drive Out through landing gear bays Fuel Pumps Out through landing gear bays Landing Gear Nose and weapons bays Oxygen system Through weapons bays 18P,S Fuel tank Remove du'ough top Inverted fuel tanks 19P,S of wing Bottom wing ports 20P,S Wing Tanks 2O6 REFERENCES FOR CHA.FrER I0 I0.I Taylor, J.W.R., Jane's All the World's Aircraft. published annually by Jane's Publishing Company, 238 City Rd., London ECID ZPU England.

10.2 Sweetman, Bill, editor, The Great Book of Modern Warp_ lanes, Salamander Books, Ltd.,New York, New York, 1897.

10.3 Taylor, R.J., Weight PredictionTechniques and Trends for Com_oosit¢ Materials Su'uctures. S.A.W.E. paper No 887, P.O. Box 66024, Los Angeles, CA 90054, 1971.

10.4 Anonymous, MilitaryHandbook 5E.

10.5 Sheinker, A. (etal) "Evaluation& Applicationof PrealIoyed Titanium P/M parts for Airframe Su'ucmres", The International Journal of Powder Metallurgy, Vol. 23, #4, 1987, pgs. 171-176.

10.6 Conversation with H.W. Smith, University of Kansas, 27 February 1990.

10.7 Niu, M.C.Y., Airframe Structural Design, Technical Book Co., 2056 West'wood Blvd., Los Angles, CA 90025, 1988.

10.8 Roskam, J., Part IV: Layout of Dcsi_ of Landing gear and Systems. R.A.E.C., Rt.

4, Box 274, Ottawa, KS, 66067, 1986.

10.9 Jacob, Tavis, Promess Re_m3rt on Monarch Wea_oons Systems Intem-ation, AE 622 University of Kansas, 4 March, 1990.

10.10 Lawson, Todd, Nozzle Desi_ and Suckdown Prediction for the Lift + Lift Cruise Supersonic STOVL fighter, AE 622 University of Kansas, 4 March, 1990.

m 10.11 Cox, Brian,et al,Prclimina_'y Stabilit-y and Control of the Lift Configm'ation, AE 622 University of Kansas, 4 March, 1990.

10.12 Roskam, J.,Part HI: Layout I)¢si_ of Cock'pit. Fuselag¢, Wing and Empennage: Cutaways and Inboard Profiles. R.A.E.C., Rt. 4, Box 274, Ottawa, KS 66067, 1986.

10.13 Hahn, A., Memo to Brian Cox, NASA STOVL group, 1990.

10.14 Zaidcl, S., Impact of Advanced Materials/Structmal Concepts on Future Weight _, S.A.W.E. paper No 1849, MacA£r, St.Louis, MO, 1988.

10.15 Witt, Borchers, and (etal),PreliminaryDesign of a Low Cost Export Fighter, Universityof Kansas, 1989.

10.16 Smith, H.W., Aerospace Materialsand Processes,University of Kansas, 1988.

2O7 1 I. SYSTEMS LAYOUT The purpose of this chapteris to document the systems layout of the Monarch. The system selections area result of the system research documented in Reference 11.2. The preliminary research was done to select the systems for the Monarch and to help in determining the layoutof these systems. The following six aircraft were investigamd: * Fairchild RepublicA-10, * GeneralDynamics F-16C, * McDonnell Douglas/BAe AV-gB, * Dassault-Breguet Rafale, * Saab JAS 39 Gripcn, * and the Eurofighter European Fighter Aircraft (_FA/_f-90).

These aircraft were considered since they have modem systems with mission requirements similar to the Monarch. From this research, the systems of the Monarch were selected.

The survivability, maintainability, and reliability of a fighter aircraft islargely dependantupon the general arrangement of the systems. The following is a list of the causes of losses of single-engine aircraft in Viemam and the Middle East;shown to exemplifytheimportance of designing survivability into the aircraft systems (Reference 11.3): 62% due to fuel system damage 18% due to pilot incapacitation 10% due to flight controls damage 7% due to engine power loss m 3% due to structural damage Much of thediscussion of the system survivability istoken from Reference 11.4. The n A work in Reference11.2was completed to determine the existence of system conflicts.

thorough conflict analysis was completedand all system conflicts were delete& This chapter willdiscussthe general layout of the Monarch systems. These systems include: II.I) landinggear, ii.2) fuelsystem, ii.3) flight controlsystem, 11.4) electrical system, ii_) environmental control system, 11.6) hydraulicsystem, avionics selection,

11.7)

and ECM selection.

11.8) 11.1 LANDING GEAR LAYOUT This section will discuss the layout of the Monarch landing gear. The material is organized as follows: 11.1.1 Gear Loads and Tire Selection 11.1.2 Su'ut Sizing and Sn'oke Length Determination 11.1.3 Landing Gear Layout 11.1.4 AircraftTip-over Criteria Methods used arc from Reference 11.1 and calculations used for this sectionare presented in Appendix 4.

11.1. I Gear Loads and Tire Selection Knowing the range of travelof the center of gravity, aircraR weight, and placement of the nose and main gear, the loads imposed on the landing gear are determined in Appendix 8. The determined loads of interest arc: * maximum static main gear load: 15,550 Ib * maximum static nose gear load: 5,400 Ib * minimum static nose gear load: 3,690 Ib * maximum dynamic nose gear load: 8,450 Ib SoR field tires were chosen for operation of the Monarch in austere rough field environments. The tire pressure on these tires is limited to below 140 psi. Since the tires chosen are designed for higher loads, the tires may bc inflatedsomewhat below the design inflation pressure for better softfield operations. The specifications of the selected tires are selectedfor the Monarch are in Table 11.1: n Table 1I.I Monarch Landing Gear Tire Specifications

Nose

Outside Diameter 22 in 31 in Width 8 in 13 in Hub Diameter 10 in 12 in Design Pressure 110 psi 135 psi Maximum Load 8,500 lb 17,200 lb Max Landing Speed 190 mph 210 mph Loaded Tire Radius 9 in 12.4 in 11.1.2 Strut Sizing and Stroke Len_h The following su'oke lengths are determined for the landing gear. It should be noted that liquid springs shock absorbers are used. A sink rate of 15 fps is used. Air Force requirements specify 10 fps. The higher sink rate was selectedbecause it is envisioned that short no-flarelandings at a steep approach angles will be employed during the service of the Monarch.

2O9 The following strut dian_u_rs and stroke lengths and are determined using landing gear load factors of 3.0 for the main gear and 11.3 for the nose gear: Nose Gear: 3.25 in diameter 16 in stroke length Main Gear:. 4.23 in diameter 11 in stroke length 11.1,3 Landing Gear Layout Nose Gear Description The nose gear layout is illustrated in Figm'e 11.1. The length of the gear designed to produce 2.5 degrees of ground incidence. It retracts forward underneath the cockpit. It is designed with 3.0 inches of trail. The gear is retracted by actuating on the drag strut.

Main Gear Description The main gear is designed with a triangulated strucULre much like that of the General Dynamics F-16 Fighting Falcon and is illustrated in Figure 11.2. This design has advantages in that: * increase in energy absorption by the tire moving laterally across the runway, * and it has a relatively large wheel stroke compared to the strut stroke.

The tire and hub section rotate 90 degrees about a line through the side strut upon retraction to lay the tire fiat in the wheel well. The oleo shock strut will act as the radius link when it is in the extended position. When the gear is down and locked, the hub is locked to keep from rotating upon landing. The gear is actuated from the drag strut which is attached to the side brace.

11.1.4 Aircraft Tip-over Criteria The tip-over angles measured for the Monarch from Figure 11.3 are: Lateral Clearance Angle: 28 deg Longitudinal Clearance Angle: 14 deg Lateral Tip-over Angle: 65 deg Longitudinal Tip-over Angle: 20 deg A takeoff analysis has been performed and verifies that the 20 degree longitudinal tip-over angle is acceptable.

11.2 FUEL SYSTEM As stated previously, the fuel system is the primary contributor to the vulnerability and survivability of an aircraft. The 'kill modes', or the types of failure, of a fuel system are (Reference 11.4): * fuel supply depletion, * in-tank fire and explosion, * void space fire and explosion, * sustained exterior fire, * and hydraulic ram.

SCAI_ 1".20 FS 180 -WL 146 WL 140 \ •',. ,,,- ../,.

WL 124 • SIDE BKACES DRAG BRACE OLEO SHOCK STRUT 16" STROKE

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Fim_c 11.1 Monarch Nose Landine Gear Layout FOLDOUT FRAME 1:20 ACTUATOR NOT SHOWN WS 41 WS 13.5 WS 45 WS 4.9 DRAG BRACE OLEO SHOCK S_ WS 0 SIDE STRUT, STATIC GROUND LIN]

I

WS 4.5 FJmn'e 11.2 Monarch Mai_ FOLDOUT FRAME FS 470 FS 483.8 ACTUATOR FS 475 I --i -- 'WI.. 124 \ WLSO Landing Gear Layout SC_ NONE FOLDOUT FRAME /.

P'WDCG w m

/.../ '-\__ _%

WS 45 N.G.

l_J 180 WS0 w Fimlrc11.3 Monarch 'r]_v i _i _ FOLDOUT FRAME _",, C_ 28 deg LONGITUDINAL TIP-OVER_ 20 AFTCG LONGITUDINAL CLFARANCE: 14 dog

O

These factors were kept in mind when laying out the fuel system of the Monarch.

11.4. The tank fuelvolumes arcas The fuelsystem layoutis shown in Figure follows: 1348 lb each wing 27.5 cu.ft. I 2450 lb forwardfuselage tank 50.0 cu.ft- / aft fuselage tank 75.0 cu.ft. / 3675 lb TOTAL 182.0 cu.ft- / 8918 Ib This accommodates the required fuel capacity of 8642 lb determined in Chapter 5.

w These fuel tanksare self-sealing and tearresistant. They are also equipped with a reticulated foam filling (porous foam panelsfolded to fell the tanks)to prevent largeullage overpressurcs following ignition of theflammable vapor (Reference11.4).As well as suppressing tankfires, the foam helpstorelieve fuelsurgingand sloshing, and may reduce the effects of hydraulic ram. Hydraulic ram istheintense pressurewaves generatedin the contained liquid by penetrators or fragments passingthroughthe liquid.

The system is set up such that there are essentially two separate systems. The forward fuselage tank is connected directly to the left wing tank, and the aft fuselage tank is connected directly to the right wing tank. The two 'separate' systems are also connected to allow for fuel management. This system allows for the complete shut-off of a damaged tank. The two systems have their own fuel pumps. These are located in the aft fuselage next to the engine. This allows for the use of a fuel suction system rather than a boost system. This is desirable since a boost system would tend to continue sending fuel through a damaged line, whereas a suction system would not be able to do this. The lift engine has its own fuel pump and line.

The Monarch is designed for single point refuelling on the underside of the left wing. It is also capable of in-flight refuelling, F-16 style, through the port in the upper fuselage behind the cockpit. Inverted Right tanks arc located within the wing fuel tanks to ensure the availability of fuel in inverted conditions. The Monarch is also equipped with a fuel management system to control center of gravity travel, a fuel jettison system through the outboard section of each wing, and a fuel indicating and ventilation system.

The engine is started with the use of a jet fuel starter. This islocated in the aft fuselage beneaththeengine and also actsas an APU. The jetfucl starter, which is essentially a smalljet engine, is started by a mechanical control from the cockpit. This releases pressurized hydraulic fluid which flowsintothejetfuel starter gearbox, starting the smallengine. This in turndrivesthe generators providing electrical power to start the main engines. The jetfuel starter is a selfsustaining system;the hydraulic accumulators are self charging after enginestart. This system requires no battery and incorporates a hand pump for hydraulic backup.

11.3 FLIGHT CONTROL SYSTEM Because maintaining aircraft stability and control is one of the most critical factors affecting safety of flight, as well as the combat survival of the aircraft and crew, much FOLDOUT FRAME 1. Fuel Pump 2. Forward Fuselage Fu¢l Tank 3. In-Flight RcfueUing Port 4. Aft Fuselage Fuel Tank 5. Wing Fucl Tank 6. Inverted Flight Tank 7. Fuel' J.¢ttison • 8. Jet Fuel Stan_

9. Undo .r_. g _fue_g Port

w m = Fuel System L Fi_n'c 11.4 =_ .

FOLDOUT FRAME _-., tout of the Monarch Aircraft attention should be given to the design of the control system to ensure that there is no unacceptable degradation of functional capabilities due to one or more component failures.

The flight control system of theMonarch is a quadruple-redundant fly-by-wire system. The flight control systemlayoutis shown inFigure 11.5. The system usesa combinationof rotaryand electrohydrostatic actuators. The leadingedge devices, the outboardailerons, and the inboard flaperons use rotary actuators; the differential stabilizer useslinear electrohydrostatic actuators. The rotaryactuators were determined to be 8 inches in lengthand the linear actuators arc required to be 2.4 inches in diameter.A uniquefeature of the Monarch's flight control system is that it has no rudder. Directional control is obtained by thrust vectoring. This isexplained in more detail in Section 9.9 For simplicity and maintainability reasons, one typeof actuator shouldbe used forallof theflight control surfaces. However, due to theproven performanceof using linear actuators in thedifferential stabilizers of other tighten, this was not done. The advantage of usingthe elvctrohydrostatic actuators is thatthey are self contained.They need not be hooked up to thehydraulic system. This increases the survivability of the aircraft since therewillnot need to be hydraulic lines to allof theactuators. The flight control system is driven by thegenerators shown in Figure 11.6.

STOVL requirements createadditional complexityin the flight control system. The aerodynamiccontrols must be Linked to thereaction control system to be used in _ansidon and hover. The reaction control systemisdiscussed in greater detail in Subsection 7.3.1.

Although it was not incorporated in the design, it was determined that it would be desirable to utilize the separate surface control system concept (Reference 11.5) for the Monarch flight control system. The following description of this concept was taken from Reference 11.5. The conventional Right control surfaces are separated into segments.

Some are driven directly by the pilot while the others are used for stability augmentation, autopilot control and attitude command applications. The servo-driven separate surface control can be used for stability augmentation functions as well as for autopilot functions.

The pilot may elect at any time to fly the airplane through the wheel while retaining furl benefit of stability augmentation.There isno feedback from the separate surfaceto the pilot.

11.4 ELECTRICAL SYSTEM The electrical system of the Monarch isshown in Hgure 11.6. The electrical system is dual redundant, powered by two 30 kVA engine drivengenerators. An electrical load analysis is shown in Figure 11.7. The phaseslisted referto themissionphasesfound in Hgure 9.1. The aircraft will still have the use of critical electrically powered components in the case that only one of the generators is operative. A 20 kVA battery is w available for backup power in the event that both generators fail. The battery will supply "N adequate power for critical equipment, such as the flight control system. Since the landing gear is designed for gravity drop extension, backup power is not required to power the hydraulic system.

The auxiliary power unit (APU) is a jet fuel starter, and performs the dual role of engine starting and backup power. The jet fuel starter is discussed in more detail in Section 11.2. No battery is required for engine startup with the jet fuel starter system.

FOLDOUT FRAME /.

I. Control Stick 2. Drive'Motor 3. Rota_/' Actuator 4. Linear Elcctrohydrostatic Actuator w Fimn'e 11.5 Flight Control System Layc FOLDOUT FRAME 3 3 t of the Monarch Ai.N-raft .

FOLDOUT FRAME I. Radar,Equipment, and Forward Avionics 2. Cockpit Controls 3. Lift 'EngineActuators 4, AuxiliaryInletActuators 5. I.ntcmal Weapons Bay Doors and Launch Mechanism 6. Aft Avionics Bay 7. Weapon Launch Mechanism 8. VentralNozzle Actuator 9. Engine Driven Generator 10. Chaff and Flare Dispenser 11. ECM and IFF Pod _2. Pitch Vae, c Actuator 13. Yaw Vane Actuator Z w = _± ; i H mu'e 11.6 Electrical System Lay, 10 12 it of the Monarch Aircra_ All of the actuation mechanisms in Figure 11.6, represented by boxes, are either elec:ro- hydrostatic or elecm3mechanical actuators and arc powered by the two generators, as arc the fli_hc control systcm actuators shown in Figure 11.5, Phase 1 2 3 4 5 6 7 8 9 10 11 4O 3O |.

> 25 " 2O

"] 15

wm Fi_m'e 11.7 Electrical Load Profile DiaeTam for the Counter Air Mission 11.5 ENVIRONMENTAL CONTROL SYSTEM The environmental control system of the Monarch, shown in Figure 11.8, uses engine bleed air. This air is piped forward from the engine compressor to the air conditioning unit and heat exchanger. The air is then used to cool the cockpit as weI1 as the avionics compartments. This system is alsoused to provide for cockpizpressm-iza6on.

FOLDOUT FRAME /" 1. To Cockpit and Forward Avionics and Radar 2. On Board Oxygen Generating System 3. Heat Exchanger and Air-Conditioning Unit 4. To Aft.Avionics Bay 5. Engine Compressor Bleed w I L = w Fi_'_ 11.8 En-,d.mmmcntal Control FOLDOUT FRAME _ , / ystem Layout of the Monarch _rt An on board oxygen generating system will be used to provide the pilot with the necessary amount of oxygen. The bulk of thesetwo systemsislocated between the forwardfuselage fuel tank and thededicated lift engine,and above theinternal weapons bay.

11.6HYDRAULIC SYSTEM The Monarch has a dual hydraulic system, both fed from the same reservoir. The hydraulic system layoutis shown in Figure 11.9. The reservoir and pumps arc located between the fuselage fueltanks and above the ammunition drum. The pumps are electrically powered by the engine drivengenerators. Hydraulicpower isused for:.

u nose gearretraction, main gearretraction, weapons bay door actuation, and gun faring.

Each of the systems has a pressure of 4,000 psi and flow rates of 40 - 50 US gallons per minute. This higher pressure willallow for smalleractuator sizes.

Hydraulic accumulators are located at the landing gear mechanisms to allow for emergency use with the hydraulic pumps disabled.

11.7 AVIONICS SELECTION An examination of current fighter aircraft was made to determine what type of equipment is required in the modem combat environment. The effectiveness of a combat aircraft is closely related to the effectiveness of its radar. Since the Monarch is required to completeboth airsuperiority and battlefield air interdiction missions, a multi-mission radar was deemed necessary.Itwas also desiredthattheradarhave day/night and all weather capability.

A systemof multi-function displays and a I-RYDwill provide the pilot with pertinent information. These willbe designedto lessenthe pilot's workload and allow him to concentrate on thetask at hand An IFF transponder willbe used foridentification purposesin combat. An air data computer and Right control computer willbe used by the Right control system. A weapons control system isrequired to provide thepilot with efficient methods of deployingweapons. Communication is_hieved with the use of UHF/VHF wansceivers and navigation isprovidedby an inertial navigation system as well as TACAN and ILS systems. The Avionicsbays of theMonarch aircraft are shown in Figure 11.6.

11.8 ECM SELECTION Susceptibility isDefined in Reference11.4as the inability of an aircraft to avoid beingdamaged in the pursuit of itsmission, and itsprobability of being hit.The bestway h_ m to decrease an aircraft's susceptibility is to make it invisible, or stealthy, to the enemy.

However, when this is not feasible, electronic counter measure devices can be used to warn of an impending attack and to provide a means to counter thatthreat.Much of the following discussion of the ECIVl devicesselected for the Monarch were takenfrom FOLDOUT FRAME /" I. N0s¢ Landing Gear Retraction Mechanism 2. Hydraulic Pump 3. Hydeaulic Reservoir 4. Main Landing G¢_ Retraction Mechanism w Fimlre 11.9 Hydraulic S, FOLDOUT FRAME <:_, tern Layout of the Monarch Aircraft Reference 11.4.

The Monarch aircraft is equipped with a radar warning receiver (RWR). This allows for the detectionof radiatingthreat elements and the accurate locationand statusof the weapon delivery systems intenton destroying the aircraft. The selection of an RWR is heavily influencedby the aircraft mission requirements. Each mission places certain requirements upon the RWR system. An RWR that is capable of meeting all of the Monarch's mission requirements, air-suporiority as wcU as battlefield air interdiction, should be selected The Monarch will use a jamming device that generates and directstransmission of a noise-like signalthat has the characteristics of radar receivernoise.Jammcrs am often used to mask or obscure the target echo. Another ECM featurethatwill be used on the Monarch are expendables. These am materials or devices designed to be ejected from an aircraft for the purpose of denying or deceiving threattracking systems for a limited period of tirnc. The Monarch will use chaff and flare dispensers at the base of the vertical tail for this purpose.

Since the Monarch aircraft has no rudder, thcr_ should be sufficient room to store this ECM equipment in the vertical tail.The ECM pods are shown in Figure 11.6. This equipment will decrease the Monarch's susceptibility to attack and the resulting damage of that attack.

w w REFERENCES FOR CHAPTER 11 11.1 Roskam, Jan, Airp_ lane Design. Part IV: Layout Design of Landing Gear and Systems. Roskam Aviation and Engineering Corporation, Ottawa, KS, 1986.

11.2 Gomer, Charles, preliminm'y_ Systems Layout of the Monarch Lift + Lift/Cruise Supersonic STOVL Fighter, University of Kansas, AE 622, March 1990.

11.3 Bonds, Ray, The Great Book of Modem Warplanes, Salamander Books, Ltd., 1987.

u 11.4 Ball, Robert E., The Fur)d_mentals of Aircraft Combat Survivability. Analysis and _, American Institute of Aeronautics and Astronautics, Inc., 1985.

11.5 Roskam, Jan, Ah'plane Flight Dynamics and Automatic Controls: Part II, Roskam Aviation and Engineering Corporation, Ottawa, KS, 1979.

u l], WEAPONS SYSTEMS INTEGRATION The purpose of this chapter is to document design work and derisions made regarding the weapons systems integration. The Monarch is designed for three different missions.

Counter Air Mission * M61 Vulcan 20ram cannon with 400 rds * 2 Short Range Air-to-Air Missiles * 2 Medium Range Air-to-Air Missiles Battlefield Air Interdiction Mission #I * M61 Vulcan 20ram cannon with 400 rds * 2 AGM-88 HARMs * 4 Mk-82 Bombs Battlefield Air Interdiction Mission #2 * M61 Vulcan 20ram cannon with 400 rds * 4 AGM-65 Mavericks * 2 Mk-82 Bombs Section 12.1 will detail the integration of the M61 Vulcan and ammunition drum.

Section 12.2 will discuss the Counter Air (CA) mission weaponry while Section 12.3 will discuss the Battlefield Air Interdiction (BAI) mission weaponry. Section 12.4 will present a description of a constructed scale model of the internal weapons bay.

12,1 INTEGRATION OF THE M61 VULCAN CANNON This section will address the integration of the M61 Vulcan cannon and the required ammunition drum and ammunition feed system. The material in this section is organized as follows: M61 Vulcan Cannon Placement 12.1.1 Ammunition Drum Placement 12.1.2 12.1.3 Structural and System Requirements 12.1.1 M61 Vulcan Cannon Placement Table 12.1 presents information available on the M61 Vulcan cannon: Table 12.1 M61 Cannon S__eifications fRef 12.D 264 _bs Uninstalled Weight: RPM Maximum Rate of Fire: 6,000 3,980 lbs at 6,000 RPM Average Recoil Force: MS0 series, 20ram Ammunition (400 rounds): 0.55 lbs Unit Amino Weight: Table 12.1 contint_¢d M61 Cannon Specifications (Ref 12.B Overall Length: 74 in Barrel Length: 53 in Maximum Diameter: 9.90 in Muzzle Diameter: 4.90 in Factors considered in placement of the cannon in the airframe include: * avoiding locations where muzzle flashes may degrade night vision of the pilot, * avoiding locationswhere the engine inlet may possibly ingestmuzzle exhaust gases, * locating the cannon with ample fuselage volume for maintenance, * and avoiding locations near vibration sensitive sensors.

For the Monarch to be an effective all-weather, day-or-night fighter, the placement of the cannon muzzle must be so thatthe muzzle flashes do not enter the director peripheral visionof the pilot. Thus, the muzzle must be located eitheron the bottom surface of the fuselage where the fuselage shieldsthe muzzle from the pilotor sufficiently aft of the pilot.

Gun exhaust gasses are highly corrosiveto the fan, compressor, and turbine blades.

Every effort should be made to assure that these gases do not enter the inletsof the engine. This may be accomplished by placing the muzzle: * behind the inlets as on the Dassault Mystere, * outboard of the inletsas on the McDonnell Douglas F-15 Eagle, * so thatthe slipstreamaround the fuselagecarries the exhaust gases away from the inlet capture area as on the Northrop F-20 Tigershark and McDonnell Douglas F/A- 18 Hornet, * or where the wing or fuselage shields the inlets from the exhaust gases as on the Mig-29 Fulcrum and General Dynamics F-16 Fighting Falcon.

To satisfy the constraints mentioned, the cannon may be placed: * in the blended area between the fuselage and the upper surface of the inlet and behind the inlet face, * or below and behind the inlet face.

The second option is chosen for the Monarch because: * the shape of the fuselage(expanding from the narrow nose/cockpit sectionto the wide aft fuselage section allows for relatively easy exposure of the muzzle, reducing the need for special failings around the gun or muzzle; * the muzzle is hidden from the view of the pilot since it is relatively far aft of the pilot and beneath the inlet, * and sufficient volume is nearby for the ammunition drum and feed system.

r _ Figure 12.1 presents this location of the cannon in the Monarch airframe. Two problems arise from this cannon location, and need to be addressed: E

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At 6,000 rounds/min, the maximum yawing moment created by the recoil force is 11,600 ft-lbs. However, 400 roundsof ammunition last only 4.1 seconds (allowing one-third second for the cannon to reachits maximum rateof tim) (Ref 12.2).

This largeof a moment isnot a likely scenario sincethe pilot ismore likely to fareseveral small bursts to conserveammo. A larger magnitude problem in the F-15, where the cannon placementis 67" outboard and the amino drum contains 940 rounds,required no rudder feedback (Ref 12.2).

Nearby accelerometers and other sensorsmay have to bc insulated from the vibrations created by thefiring cannon. This proved to be the cause of two F-16 accidents in the summer of 1979. Cannon vibrations were sending false readings to theflight control computer (Ref 12.2).

12.1.2 Ammunition Drum Placement w The missionspecifications for theMonarch stipulate 400 20ram rounds. This amount of ammo requires 5.2 cubicfeet of volume for thedrum. The dimensions for the drum arc: Diameter: 18" = Length: 35" w Given the gun placement, the ammunition drum may be located either:.

* behind the cannon as on the Grumman F-14 Tomcat (see Figure 12.2) and Fairchild A-10 Thunderbolt, below the inlet, and between the internal weapons bay and outer fuselage, * or above the internal weapons bay and between the bifurcated inlets similar to the General Dynamics F-16 Fighting Falcon, see Figure 12.3).

The first option is desirable because of the ease of access to the drum for maintenance and reloading. It also does not sacrifice any fuel volume in the center of the aircraft. However, there is insufficient volume available for clearances (less than 2" of clearance laterally), the weight of the amino is displaced from the aircraft centerline, and complicated twists arc rextuired in the feed system.

The second optionis chosen forthe Monarch primarily for simplicity. The Monarch ammunition drum and feed system is shown in Figure 12.4on the previous page. The system ismuch likethatused on the F-16, sec Figure 12.2. The drum islocated above the weapons bay and below the upper fuselage surface. Access to the drum for maintenance reasonsisthrough accesspanels on the upper fuselage.

The feed system useslinkless ammunition and fits throughthe 6" clearance between the corner of the weapons bay and theinlet.New rounds of ammunition from the drum feed tothe cannon. The used casings then arc fed above theweapons bay and underneath the ammo drum to the starboard sidewhere they are fed back intothe amino drum.

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,= _Z i r L w Z ¢,/3 ORIGINAL PAGE IS OF POOR QUALITY w = w w the new rounds arc loaded through the starboard loading access panel.

Reloading of new carnidgvs and unloading of used casings is accomplished through an access door on the starboard siclc of the fuselage underneath the inlet that is w approximately 1 foot square. This low location will allow for easy reloading by armorers since they do not need to climb on the aircraft or use ladders. However, simultaneous loading of ammunition and Sparrow missiles into the internal missile bay will bc haml_red.

This is ratherunavoidable due the locationof the bay. The problem could bc avoided ff the cannon was nose mounted. The problem may bc reduced by simultaneously reloading ammo and the port side ordnance before loading the starboard ordnance.

12.1.3 Structural and System Re.__t)ircments Structural Requirements The Vulcan cannon will mount on a frame at FS 350. This frame is currcntiy designed for severalother uses (see Chapter 7). Other gun mounts where the barrelwill W need support arc at FS 283 and FS 313. The ammunition drum will also mount to the bulkhead at FS 350 above the internal weapons bay. Cut outs will have to be provided for = the muzzle at FS 276 (with appropriatecooling vents), for the reloading access panel, and for the amino drum and cannon access panels.

+ + System Requirements w The Vulcan and feed system may bc powered electrically, hydraulically, or both.

The cannon is triggeredelccu'icaUy, by a signalfrom the cockpit (Ref 12.2).

= _ 12.2 COUNTER AIR MISSION WEAPONS INTEGRATION This section will present the deraildesign and decision made regarding the weapons integration for the Counter Air mission. The materialin this chapmr is arranged as follows: r .

12.2.1 Short Range Missile Inmgradon 12.2.2 Medium Range Missile Integration 12.2.3 Sn'uctural and System Requirements The Counter Air mission uses the following weapons: * 2 Short Range Air-to-AirMissiles * 2 Medium Range Air-to-Air Missiles The A._-9 Sidewinder and AIM-7 Sparrow arc curmndy in the US inventory as short range and medium range air-to-air missiles, respectively.The ASRAAM and AMRAAM projects, replacements for the Sidewinder and Sparrow, arc expodencing difficulties. Since the A.]_-9 and AIM-7 missiles arc largerthan the ASRAAM and AMRAAM, the weapons integration will bc designed for the larger missiles, assuming the smaller missiles can bc integrated easily in the fum.rc. Table 12.2 presents available data for the Counter Air mission weapons: Table 12.2 Counter Air Weaponry_Data (Ref 12.1_

AMRAAM ArM-TF AIMcg2 An_-9L ASRAAM

u Guidance Radar Semi- IR IR fIR Method Active Radar Range (miles) >100 24 6 10 11.9 ..==, Launch Method Eject Eject Rail Rail Rail or Rail w Weight 0bs) 327 614 170 191 150 = Overall Length (in) 144.0 146.0 113.0 115.0 98.4 Body Diameter (in) 6.8 8.0 5.0 5.0 5.9 w Fin Span (in) 25.0 40.0 25.0 24.0 17.7 12.2.1 SHORT RANGE MISSILE INTEGRATION It was originally intended to carry the Short Range Missiles internally. However, after further consulting (Ref 12.3), it was decided to mount the AIM-gLs on wingtip launchers because: * pre-launch target acquisition is rebuild.

* the wingtip launchers provide the AIM-gL with a larger field of view, allowing better acquisition of targets. The proximityof the fuselage severely limitedthe field of view ifinternal storage was employed.

* wingtip launchers increase the effective aspect ratio of the wing, thus increasing thewing efficiency and reducinginduceddrag (Ref 12.4).

* reliability and simplicity of the system increase the effectiveness of the weapon and reducethe cost of both theweapon and aircraft.

The disadvantages of mounting the Sidewinder missiles on thewingtips are: * a decreasein stealth, * an increase in parasite drag, * and an increase in rolling moment of inertia.

Figure 12.5illustrates thewingtiplauncher and Figure 12.6illustrates a schematic diagram of themissile-restraint devicewithinthewingtiplaunchrail.The railmust guide the missile duringlaunch,yetitmust hold the missile in place during+9/-3g maneuvers and retain themissile in the case of an accidental rocketmotor ignition. With the locking mechanism in place, the missile isnot allowed to move until the missile is selected and the lockingmechanism israised.The blocking mechanism spring isdesignedto retain the missile in a low acceleration environmentso thatupon motor ignition the missile lug will 150 D t_J z u z m = 5O w i ul uJ z u .J =__ -50 tJ t- - 100 m - 150 - 2O0 - 25O

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12,2,2 MEDIUM RANQE MISSILE INTEGRATION Internal storage of the medium range missiles is specified in the mission requirements. The Sparrow may either be ejected or rail launched, it is decided to eject the missile since the ejection mechanism requires less volume than a retractable rail launcher.

Due to the large span of the AIM-7F (40 inches), the fins are folded. Figure 12.7 presents the fin folding scheme of the AIM-7F. The fin fold layout was determined considering: * all four fins should be the same and have only one hingeline to reduce the cost.

* the minimum volume is created by making the missile "square" when stored.

* the fins may not strike each other when they deploy.

_m w Launch Sequence The medium range missile may incorporate systems similar to those used on the AIM-7M in the future. The AIM-7M uses the AIM-7F airframe but incorporates a new digital guidance section. It is designed for improved capability in look-down and ECM environments. The AIM-7M also incorporates a LTE/BIT. During the Launch-To-Eject (LTE) cycle a Built In Test (BIT) is conducted. If the LTE/BIT detects a missile failure, the missile will not be launched. The LTE cycle typically requires 1 to 2 seconds depending primarily on the time required for gyroscope run-up which typically takes 0.75 to 1.5 seconds (Ref 12.5). Figure 12.8 presents a typical LTE cycle. However, in the case of the internal weapons bay, the weapons bay doors open when the trigger is pulled.

Since the pilot currently experiences a 2 second delay between trigger and launch, it is believed that the door actuators may be sized to open at more than 45 deg/sec. The BIT m will have an additional "doors open" test before ejection. Upon jettison of the missile, separation of the umbilical, and predetermined linear acceleration, the fins deploy and the rocket motor ignites. The missile then undergoes a preprogrammed maneuver until it clears the aircraft.

Ejector Design For safe, reliable and effective use of the AIM-7F, the missile must be ejected clear of the aircraft. To achieve clean separation, major concerns that need to be tested and simulated are (Ref 12.5): * that the missile will not strike the launch aircraft during powered (motor-fire) or jettison (no motor-fire), * that the missile will not strike the aircraft as a result of a failure during separation, * and that the missile rocket blast will not _lversely affect the launch aircraft.

Key parameters for design of the required ejector are the linear velocity and angular velocity imparted to the missile at the end of the ejection stroke. Typical linear velocities are 18-20 ft/sec and typical nose-down angular rates are from 0-30 deg/sec depending on ].w S t I v v..._

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the weight, moment of inertia, and installation of the missile (Ref 12.5).

The limiting factor on linear velocity is the structural limit of the missile. To achieve 20 ft/sec linear velocity, average accelerations of 10-20 g's and peak accelerations of 30g's are experienced by the missile. The angular rate is limited by the capabilities of the seekerhead stabilization loop (Ref 12.5). Because of the ejector location (inside the fuselage as opposed to on the fuselage surface, as is the case on both the F-15 and F/A-18), it is assumed that the ejectors will have to be powerful and deliver high linear I and angular velocities. To determine the correct velocities, extensive simulation, tunnel modeling, and full-scale flight testing are required (Ref 12.5).

_m A typical ejector isillustrated in Figure 12.9.The missile is mechanically attached to the ejector. Separate ejector feetare used to eject themissile.Upon ignition of the pyrotechnic cartridges in the breech(ffone fails to ignite, one willignite theother), the exhaustgases release the mechanical linksbetween the missile and the ejector as the required pressure buildswithinthe ejector cylinders. At the end of the ejector stroke, the gases are ventedout of thepistons and simple springs retract thepistonintothe cylinders.

The required angular rateand linear velocity arecontrolled by a gas flow control valve between thebreech and each ejector cylinder (Ref 12.5).The pyrotechnic cartridges may be replaced in the ejector rack from thesidewithoutremoving themissile.

Missile Bay Design Given the allotted volume and required missile and ejector rack volumes, the minimum volume for the internal missile bay is: Length: 156" Width: 48" Depth: 22" The design of the internal missile bay and ejector areshown in Figure 12.10. Each door operates independently and consists of two panels. The inboardpanelishingednear the aircraft centerline and two rotaryactuators rotates each door aboutthehinge. The inboardedge of the outboardpanelishingedtothe outer edge of theinboardpanel. The outer edge of theoutboardpanel follows two lateral tracks on the forward and aftends of themissile bay. When opening, theinboardpanelwillrotate down 90 degreesand the outerpanel willrotate -90 degrees while traversing inboard. This door design was chosen because: * the doors are opened only briefly and the degradation in handling qualities and increase in drag is not as severe as if the doors were opened for a longer period of time. The increase in airplane drag coefficient from the opened weapons bay and doors at Mach 1.6, 30,000ft is estimated to be 5 drag counts, an increase of 1.2% in zerolift drag. See Appendix 5 for the calculations.

* thedoor designisrelatively simple, only four moving parts: two door panels and two actuators, * thedoorsopen to thecenter so that armorers have unobstructed access to the weapons bay,

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w ORIGINAL PAGE IS OF POOR QUALITY * and the division of the weapons bay along the centerline allows for the addition of a keel beam along the fuselage and should aid in carrying the loads around the smaller cutouts in thefuselage.

The clearance angles between the missile bay and the missile are: Forward: 28 deg Aft: 16 deg Inboard: 5 deg Outboard: 8 deg Reference 12.4 suggests 10 degreesof clearance both laterally and longitudinally.

Since themissile isrelatively guidedthough the bay ratherthandropped,thelaterally clearances areassumed to be acceptable. The largelongitudinal clearances allow for door tracks and otherstructure.

12.2.3 Structural and System Requirements Structural Requirements The following structural requirements are needed for the integration of the Counter Air mission weaponry: Attachments are required on the wingtips for the launch rails. The internal weapons bay outer walls are located at: WS +/-24 FS 250 FS 356 WL 124 WL 146 The forward and aft walls of the weapons bay must provide for door track attachments. The beam along the centerline must provide for actuator attachment and may be no more than two inches wide at the bottom.

w The top of theweapons bay needs extrastructure near the ejector pistonsatFS 293 and FS 350 at WS +/-13.5.

System Requirements All that is required for the wingtip launchers is electric power and trigger signal to the launchrail.The internal weapons bay requires four rotary actuators, either electric or hydraulic.The ejector rack and missile reqttire electric power and signaling.

12.3 BATILEFIELD AIR INTERDICTION MISSION WEAPONS INTEGRATION This section discusses the integration of the weapons required for the two Battlefield Air Interdiction missions. The two mission loadings are as foUows: * BAI #1:2 AGM-88 HARM_ and 4 Mk-82's * BAI #2:4 AGM-65 Mavericks and 2 Mk-82's The material in this chapter is arranged as follows: 12.3.1 Battlefield Air Interdiction Mission Weapons Arrangement 12.3.2 Structural and Systems Requirements 12.3.1 Battlefield Air Interdiction Mission Wea__ns Arrangement Two primary factors are considered in determining the placement of the BAI weapons. Due to their large size and the lack of a supercruise requirement, it is decided to carry BAI weapons externally. This arrangement also allows simultaneous carriage of the counter air weaponry (2 AIM-9Ls and 2 AIM-7Fs). If required, two auxiliary fuel tanks may be fitted into the internal weapons bay which add approximately 35 cubic feet or 1,700 lbs of additional fuel (see Appendix 5 for the calculations).

Table 12.3 presents available data on the BAI ordnance: Table 12.3 BAI Ordnance Specifications fRef 12.1) AGM-88A AGM-65 Mk-82 Mk-82 HARM Maverick Slick fzllgk¢_ Guidance Method Radar TV None None ]JR Laser 13 miles N/A N/A N/A Range w Launch Method Ejection Rail Drop Drop 807 463 521 560 Weight (lbs) Overall Length (in) 164.0 98.0 87.0 88.5 10.0 11.8 10.8 10.8 Body Diameter (in) Fin Span (in) 44.0 28.3 16.0 16.0 Due to the largo weight of the ordnance, it is desirable to keep the weapons as far inboard as possible in order to reduce asymmetric loads during hover and the aircraft rolling moment of inertia. The considerations that limit the placement of the weapons are: * conformal fuselage mounting is rejected because of the limited space available between the weapons bay doors, lift engine, and landing gear doors. Also, during hover, the fountain core that is created impinges on the fuselage at approximately FS 400 would heat up the stores.

* the HARMs and Mavericks are considered to be "high value" stores when compared to the gravity bombs. It is undesirable to have to require the jettison of "high value" stores in order to reduce an asymmetric load for balance in hover.

Thus the HARMs and Mavericks are carried inboard of the Mk-82's.

For BAI mission #1, the HARMs are carried on an inboard wing pylon. The tail of the missile determines both the lateral and longitudinal placement of the missile in that: w * the tall must be forward of the wailing edge devices, * and the tail must be able to clear the main gear doors and the deflected gear upon landing.

w The vertical placement of the HARM is limited primarily by clearance of the leading edge devices.

This places the inboard pylon at WS 85 and the center of gravity of the HARM at: FS 400 WS 85 WL 130 The Mk-82's are carried on an outboardwing pylon. They areplaced outboard sufficient for clearance of the HARMs. The Mk-82's are mounted on a twin stores ejector.

Figure 12.11presents approximate lateral clearances of the twin stores ejector. Figure 12.12presents the current layoutof the pylonsfor BAI mission #1.

For BAI mission#2, two Mavericks arecarried on the inboardwing pylon. To allow foradequate growth,lateral space is allowedfora Maverick triple raillauncher.

The clearances are illustrated in Figure 12.11. Another constraint applicable to the Maverick carriage isthatthe seeker head of the Maverick has a 5 degreehalf-angie cone of vision(Ref 12.2).These angles are demonstrated in Figure 12.13on the layoutof the weapons for BAI mission#2.

The Mk-82's during BAI #2 are mounted on single ejector racks on the outboard pylon. The outboard pylon is placed such that allowable clearances are made for simultaneous carriage of 3 Mavericksand 3 Mk-g2's on each wing. The approximate lateral clearance of the Mk-82's on a triple ejector rack isillu_ in Figure 12.11. The outboardpylon islocated atWS 135. Figure 12.14 presems the fully loadedMonarch with 6 Mavericksand 6 Mk-82's. Of particular interest are the lateral tip over clearance angles.

As seen in Figure 12.14, allstores are satisfactorily insidetheline S inches above a 5 degreeanglelinefrom the main gear contact point(Ref 12.4).

I- Scale: 1:20 w Twin Stores Carriage TripleEjectorRack (Mk-82s) (Mk-82s) -._ 6 "I_'-- /, .',.

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12.3.2 Structural and Systems Requirements Structural Requirements The longitudinal location of the hard points is determined by those in the smacmres group. They will be slightly aft of the front spar. An additional small attachment will need to be provided near the rear spar for attachment of the pylon.

System Requirements The only systems required by the Battlefield Air Interdiction mission ordinance pylons are electric power and signalling to each of the wing pylons. Other possible requirements which have not yet been discussed are the addition of fuel lines to the internal weapons bay and the wing pylons to allow the carriage of auxiliary fuel tanks.

12.4 INTERNAL WEAPONS BAY MODEL DESCRIPTION A model demonstrating the feasibility of the internal weapons bay has been constructed. This model will be transported to the USRA conference in June 1990.

The scale of the model is one-fifth scale. One-tenth scale is too small for intricate pieces and one-halfscale is too large for the allotted conference display area.

With the close proximity of the cannon and ammunition drum to the weapons bay the model also includes the M61 Vulcan cannon and ammunition drum installation.

The model is constructed of: * bass and balsa woods in missile fins, and the cannon, * plywood in the structural flame and doors, * metal in the moving parts of the ejector, door tracks, and hinges, * cardboard for the flexiblesurfacessuch as fuselage, and amino feed system * PVC pipe for the ejector pistons and ammunition drum.

The approximate model size in 1/5 scale is: Length: 32" Width: 18" Depth: 13" Due to the proximity of the inlets and auxiliary inlets, they are incorporated into the model.

Figure 12.15 shows photographs of the completed model.

PhotDgraph pending completion of model Figure 12.15 Internal W, apons Bay Model Photo m'aph w REFERENCES FOR CHAPTER 12 12.1 Roskam, Jan, Airplane Desi_ Part IV: Layout Desi_ of Landing Gear and System_, Roskam Aviation and Engineering Corporation, Ottawa, Kansas, 1989.

12.2 Fitzsimons, Bernard; de Ste. Croiz, Philip; Bonds, Ray; and Hall, Tony, eds., The Great Book of Modem Warplanes, Portland House, New York, 1987.

12.3 Hahn, Andy, NASA Ames Powered Lift Engineer, personal conversation, February, 8, 1990.

12.4 Raymer, Daniel P., Aircraft Desi_: A Conceptual Approach, American Institute of Aeronautics and Astronautics, Washington D.C., 1989.

12.5 Eichblatt, Emil J. Jr., ed., Test and Evaluation of th_ Tactical Missile, Volume 119 of Progess in Astronautics and Aeronautics. American Institute of Aeronautics and Astronautics, Washington D.C., 1989.

13. LIFE CYCLE COST ANALYSIS The purpose of this chapter is to present the results of the cost analysis of the Monarch fighter program.

The methodology of Chapters 3, 4, 6 and 7 of Reference 13.1 were used to determine the following costs for the Monarch fighterprogram: 13.1 Research, Development, Test and Evaluation Cost 13.2 Acquisition Cost 13.3 Operating Cost 13.4 Life Cycle Cost The life cycle cost ('LCC) is made up of the research,development, testand evaluation (RDTE) cost, the program acquisition cost,the program operating cost and the disposal cost. These four components of the life cycle cost arc incurredduring the six phases of the aircraft life cycle as shown in Figure 13.1. This figurealso illustrates the percentage of the life cycle cost thatis locked in during each phase of the aircraft life.

The Monarch is currentlythrough Phase I and Phase 2 of the aircraft life cycle. This indicatesthat 85% of the life cycle cost of the Monarch fighter program is locked in. By being aware of the implications of the data presented in Figure 13.1, the design team has insightinto the influencethatdecisions made early in the design process have on the life cycle cost of the aircraft.

It is assumed that the Monarch fighter will bc in operation for 25 years beginning in the year 2005. A baseline production run of 500 airplanes and an annual utilization of 325 flight hours per airplane per year are assumed. The costs arc based on anticipated 2005 rates and US dollar value.

13.1 RESEARCH. DEVELOPMENT. TEST AND EVALUATION COST The following values were required for the determination of the RDTE cost of the Monarch fighter program: Takeoff weight = 31,336 lbs Maximum Velocity = 794 keas Number of Airplanes Built for the RDTE Phase = 10 Difficulty Factor for the Monarch fighter program = 2.0 (2.0 assumes an aggressive use of new technology) CAD Experience Factor = 0.8 (0.8 assumes CAD experience) Engineering Manhour Rate = $105.00 (reflects a 50% increaseover non-securityrate) Cost EscalationFactor = 3.1 (estimated for 2005) m

J

L_

J

i B Number of Conventional Engines Per Airplane = 1 Number of Static Test Airplanes ffi 2 Main Engine Takeoff Thrust = 35,573 lbs STOVL Equipment: Lift Engine ffi $1.027 million Lift Engine Nozzle = $154,080 1-D Vectoring Ventral Nozzles ffi $222,400 2-D Vectoring Main Nozzle = $444,800 Material Correction Factor ffi 2.5 (2.5 assumes construction with conventional composite materials) RDTE Production Rate = .35 airplanes/month Manufacturing Manhour Rate = $68.00 (reflects a 50% increase over non-security rate) ToolingManhour Rate = $83.00 (reflects a 50% increase over non-security rate) Stealth Factor --1.0 (I.0assumes no designed stealth features) Test Facilities Cost Adjustment Factor = 0.2 (0.2assumes extensive test facilities arerequired) Percentage of Profit on the RDTE Phase = 10% Financing Cost Factor = 0.13 (0.13 assumes a 13% interest rate on the financing) The value for the STOVL lift engine was determined by increasing the value of a conventional engine of equal thrust (12,105 lbs) by 20%. This increase was due to the advanced technology required to produce this engine. The lift engine is estimated to weigh 480 pounds, thus resulting in an installed thrust-to-weight ratio of 25. The lift engine nozzle was estimated to be 15% of the lift engine value. The ventral no-,]es were estimated to be 10% of the main engine value. The 2-D main nozzle was estimated to be 20% of the main engine value. These estimated values were obtained from Reference 53.

The total RDTE cost for the Monarch fighter program was determined to be 3.716 billion dollars.

13.2 A_-'QUISITION COST The acquisition cost is the cost that the government or taxpayers pay for the total number of airplanes in the program. The acquisition cost is the sum of the manufacturing cost and the profit made by the manufacturer. The following values were used as input in determining the acquisition cost of the Monarch fighter program: Number of Airplanes Built to Production = 500 Manufactm'ing Rate of Production Airplanes = 10/month AirplaneOperatingCost Per Fright Hour = $I0,146/hr (from Section 13.3) Test Flight Hours beforeDelivery = 20 Overhead Factor= 4.0 Manufacturing Finance Factor --- 0.13 Manufacturing Profit = 0.I The acquisition cost for the Monarch fighterprogram is determined to b¢ 12.206 billion dollars. Figure 13.2 illustrates the effectof the numbvr of airplanes produced on the totalprogram acquisition cost.

The average estimated price (AEP) per fighter is dctcrmine, d by summing thc RDTE and the acquisition costs and dividingby the number of airplanes produced. With 500 airplanesproduced, the average cstimatcd cost of the Monarch fighter is 32.6 million dollars. Figure 13.3 illustracs the AE,P of the Monarch fighteras a function of the number of airplanesproduced. Note that for production rims greaterthan about 600 airplanes, no significant decrease in AEP is experienced for a reasonable increasein the number of airplanes produced.

I Calculated for C I i_imat,_ 2005 I.SD 20 I C O I I I _3 u I ! 0 i I !

.2 10 ¸ I , w 0 100 200 300 400 5006007008009001000 Number of Production Airplanes Fi_mzm 13.2 Effect of Numbe1" of Airp_ I,_es Produced on the Acquisition Cost for the Monarch Fichrar PrO_m'_m

i,!o_

0 100 200 300 400 500 600 700 800 900 1000

Number of Production Airplanes

Fi_m_ 13.3 Effect of N-tuber of Ah'planes Produced on the Average Estlmate_ Price of the Monaz_h Fighter 13.3 OPERATING COST The program operating cost is the total amount of money that will be required to operate a specified number of airplanes, flying a certain number of hours per year for a set number of years.

w The following valueswere used to determinethe total operating cost of the Monarch fighter program: MissionFud Weight = 8642 Ibs Fuel Price = $1.95/gal (estimated 2005 price) Fuel Density = 6.55 Ibs/gal (IP-4) Annual Utilization = 325 fright hours Average Mission Time = .80 his (air-to-air mission) Number of Airplanes Built to Production = 500 l Annual Loss Rate Per 100,000 Flight Hours ffi 7 Number of Years in Active Duty = 25 Number of Crew Members ffi 1 Crew Ratio Per Airplane = 1.1 Ah'crew Basic Pay = $34,000 Aircrew Incentive Pay = $500/month Aircrew Re-up Bonus = $14,000 Maintenance Man Hours Per Flight Hour = 15 Cost Escalation Factor = 3.1 Airplanes Used by the Reserves Factor = 0.10 Indirect Personnel Cost Factor = 0.2 Spare Part Cost Factor = 0.18 I_pot Cost Factor = 0.16 The total operating cost for the Monarch fighter program is determined to be 28.88 billion doUars.

The operating cost per flight hour is determined by dividing the total program operating cost by the number of airplanes in service, the number of years the airplane is in active service and the number of hours each airplane is flown annually.

Therefore the operating cost per flight hour for the Monarch fighter is determined to be $10,146. Figure 13.4 illustrates the effect of the number of airplanes produced on the operating cost per flighthour for the Monarch. Note that when conducting thistrade study,the cost of program indirect personnel, consumable materials, depot maintenance and misccUaneous accrualsarc held at the baselineproduction number values. For the baseline production of 500 airplanesthese costs have the following values: Indirect Personnel = 5.776 billion dollars Consumable Materials = 286 million dollars Depot Maintenance = 4.621 billion dollars Miscellaneous Accruals = 1.144 billion dollars 13.4 LIFE CYCLE COST The life cycle cost (LCC) represents the total amount of money spent on the airplane program. The life cycle cost is broken down into the following components: 1. Research, development, test and evaluation cost 2. Acquisition cost 3. Program operating cost 4. Disposal cost of the airplanes Values for items 1 through 3 have been computed.

No accurate method exists for determining the cost of disposal. Reference 13.1 suggests that the disposal cost is 1% of the program life cycle cost. Figure 13.5 shows the breakdown of life cycle cost for the Monarch fighter program.

i

\

1 I. ,I i i

90000--

eOOOOJ 1 Calculated for B_zmted 2005 LBD

i

I -1, I

o 70000

BueiJ_e _ of I t Production Ran= = 500

*_ 60000

!

i

50000

l I , _, 40000' I

_ 3oooo ' i i

I I

_ 20000- - _ "_.T I i t

t I

I I I

1oo00

I f i i ! = t i

i

= i i

!

o 0

I i i' i

100 200 300 400 500 600 700 800 9OO IOO0

Number of Production Airplanes

Fimu'e 13.4 Effect of Number of A_lanes Produced on the .Opera_ringCost Per F].ight Hour for the Monarch Hghter

(_.ox)

(o.27.)

.456 _ 3.716 Btl]lon T,....

_uo= (83.27.) -_

_kxlu_t.ion (27.67.)

12.608 Billion 28.880 BiUJon Fi_m_e 13.5 Life C'ycle Cost Breakdown of the Monarch Fighter Pro__-s,"n 26O REFERENCES FOR CHAPTER 13 13.1 Roskam, Jan, Airplane Desim_: Part VIII. Airplane Cost Estimation: Desire1.

Development. Manufacturing and .Operating, Roskam Aviation and Engineering Corporation, Ottawa, KS, 1989.

13.2 Personal Consultation, Dr. S. Farokhi, Universityof Kansas, May I, 1990.

14.CONCLUSIONS AND RECOMMENDATIONS 14.1CONCLUSIONS

I)

The results of the Phase I studyof threesupersonic STOVL concepts indicate thatthe configuration with the lift+lift/cruise engine cycle suffered the least penalties for the short takeoff and vertical landing capability and was the concept with the most readily available technology for a Technology Availability Date of 1995.

2)

Based on thePhase I studyresults, thelift+li_cruise configuration was selected for Phase IIdesignwork. This work consisted of more detailed configuration design and concentrated on some of the $TOVL aspects of the airvr_.

3)

The aircraft has a wide,fiat fuselage sections due to the requirements for: * the lift engine, * the internal weapons bay, * the large landing gear tire sizes, * the ventral nozzles for the cruise engine, and * shaping considerations for a favorable area rule distribution.

4)

The aircraft weights were estimated using empirical equations based on statistical data and actual weights taken from operational fighters. The aircraft is balanced in hover and has acceptable inflight center of gravity (ravel.

5)

The aircraft achieves powered lift with a lift engineand two ventral nozzleson the main engine. Three posts allowed the in ground effect suckdown to be reduced to 10 percent (this represents a 15 percent reduction when compared to the Phase I li_f-t+lift/cruise configuration which had a two posts).

6)

The thrust vectoring capabilities of the cruise engine nozzle allowfor: * enhancedmaneuveringat high anglesof attack, * removal of the rudder, * and reduction in thevertical tail size.

The reduced vertical tail size aided the area rule distribution of the aircraft to very favorably match that of the Sears-Haack shape.

7)

The aircraft, without rotating for hover, can lift off in 238 ft for the design counter air mission and in less than 800 ft with an overload mission (the overload mission consists of 5,000 lbs of more ordnance than the counter air mission).

8)

The aircraft has a high level of performance throughout its flight envelopeand compares favorably to the operational aircraft of the United States and SovietUnion.

The aircraft has growth potential in that it can perform typical NATO missions with acceptable range and ordnance carrying capability. Being a STOVL configured, the w aircraft also has the ability to perform unconventional two stage missions, possibly allowingfor increased sorties per day.

w The aircraft can be trimmed at all investigated flightconditions. For three flight

9)

conditions, a digital stability augmentation system was employed for the longitudinal, lateral, and directional axis to meet Level 1 handling qualifies. The augmented aircraft is not prone to inertia coupling at the three flight conditions investigated.

An aircraft structural layout was completed with structural synergism as a key 10) priority.The materials for the aircraft were selected consideringtheirresulting weight in an aircraft application, their cost and durability, and their ease of repair in a battlefield scenario. The manufacturing process and breakdown of the aircraft was preliminarily investigated along with itsthe accessibility and maintainability considerations.

11) The system layout for the aircraft was complete and allsystem conflicts were The eliminated by using a combination of three view and ghost view system layouts.

avionics and electronic counter measures were selectedfor the aircraft.

The medium range missileswere successfully mounted insidethe aircraft. The 12) integration and launch mechanism of the internal missileswas verified by building a 1/5" scaled model of the layout. The short range missilesare wing tip mounted and the battlefield air interdiction mission ordnance are carried on wing pylons.

RECOMMENDATIONS 4,2 The aircraft cannot meet the Level 1 specification for roll performance in all flight

I)

conditions with only aileron deflection. Two approaches should be investigated to achieve the required roll performance. First, the effect on the roll performance of the deflection of the leading edge flap should be calculated. Leading edge deflection may also aid in reducing unfavorable aeroelastic effects. Second, the effect on the roll performance of the allmoving stabilator should be calculated.

Hot Gas Reingcstion (HGR) may be a problem during hover close to the ground.

2)

Since HGR is very configuration dependent and is difficult to predict, wind tunnel tests should bc performed. Nevertheless, possiblesolutionswere investigated to alleviate the HGR that can be implemented without major design changes to the aircraft. The possibleground erosion due to the jet exhaust of the lift and main engine should also be investigated with experimental techniques.

The effect of the cannon firing creating adverse yawing moments needs to be studied

3)

further to determine ff rudder feedback is required. Also, the effect of cannon vibrations in the airfl'ame should be studied to determine ff aircraft sensors need to be isolated.

The bending and torsionof the internal weapons bay doors while deflectedat a high

4)

dynamic pressure need to be investigated.The actuators for the doors also need to be sized to effectively open the door at high dynamic pressures.

The aircraft will require a ride qualityaugmentation system to allow the aircraft to

5)

successfully perform low level ground attack missions.

APPENDL 1

The purpose of this appendix is to document the weight and balance calculations presented in Chapter 5. The weight and balance was done using a spreadsheet.

Appendix 1.1 def'mes the symbols and summarizes the equations used for the spreadsheet.

Appendix 1.2 shows the spreadsheet used to calculate the weight and balance.

Avvendix 1: Table of Contents page w w CLASS II WEIGHT SPREADSHEET SYMBOLS Svmbo!

Parameter Unit S w Wing area ft 2 m A w Wing aspect ratio ....

D LMLEw Wing sweep angle deg Lm w Wing taper ratio ....

(t/c)_m_w Wing thickness ratio max ....

cbar w Wing mgc ft 1 f Length of fuselage ft w h f Height of fuselage ft Horizontal tail area ft 2 S h b h Horizontal tail span ft cbar h Horizontal mgc ft i h Distance from wing c/4 to tail cm/4 ft t r h Horizontal tail max root thickness ft Distance from the vertical tail root to where the z h horizontal tail is mounted on the fuselage.

fuselage mounted horizontal tails, z_h = 0. ft A v Vertical tail aspect ratio ....

b v Vertical tail span ft Vertical tail area ft 2 S v 1 v Distance from wing c/4 vert. tail Cv/4 ft Lmv Vertical tail taper ratio ....

Vertical tail sweep angle c/4 deg LM_c/4_v Rudder area ft 2 S r Canard area ft 2 S c b c Canard span ft t r c Canard max root thickness ft cbar c Canard mgc ft 1 c Distance from wing c/4 to canard cJ4 ft w W TO Takeoff Weight Ibs W E Empty Weight lbs D W F Fuel Weight lbs W el Engine 1 weight ibs W e2 Engine_2 weight Ibs W crew Crew weight Ibs GW Gross design weight Ibs W wtr Weight of water for injection Ibs W iae Weight of instruments and avionics ibs Weight of gun and launcher Ibs W_glw Wets Miscellaneous weight Ibs K fcf Fixed Equipment - flight control sys ....

K ec Power Plant - engine controls ....

Structural - landing gear - main wing ....

K_g_r K w Structural - wing ....

K-inl Structural - fuselage ....

Kd Power Plant - air induction system ....

w Km Power Plant - air induction system ....

Power Plant - fuel sys - self sealing lbs/gal K_fsp N inl Power Plant - air induction system ....

N e Power Plant - engine controls ....

Number of crew ....

N cr n ult Ultimate load in g's ....

D MH Maximum Mach speed at sea level qbar_D Design dive dynamic pressure ft Ld Power Plant - air induction system ft 2 A inl Power Plant - air induction system P 2 Power Plant - air induction system psi _mum Weight of main gear to nose gear W_mg/W_g Mission fuel fraction M ff _mmm Ratio of GW to takeoff weight GW/T TO Ibs Put weight of weapons in spreadsheet Weapons

A_ng B_ng C_ng D_ng PART V CONSTANTS FOR LANDING GEAR PAGE 82 NO UNITS A_mg B_mg C_mg D_mg w J WEIGHT SPREADSHEET GUIDE This guide has equations developed into a spreadsheet for w determining class II weights for supersonic fighter attack aircraft.

ProGedure for sDr@_dsheet we_qht _naly_is: I.) Follow equations in this guide and enter parameters for particular airplanes at the TOP of the spreadsheet ONLY!

2.) Page down in spreadsheet to see how weights have changed.

3.)

Split Screen and iterate takeoff weight 4.) Enter F.S., B.L., W.L., in lower portion of spreadsheet.

weights will have automatically transferred.

NOTES: Enter all areas in ft.

Enter all angles in degrees.

Enter all weights in ibs.

Enter GW/WTo fraction in TOP of spreadsheet Enter WmJWg at TOP of spreadsheet Enter M, at TOP of spreadsheet Enter WpL at TOP of spreadsheet Enter landing gear option (I or 2) at TOP of spreadsheet Option I: General gear equation Option 2: Main gear on wing

STRUCTURAL WEIGHT (lbs)

Wstruct == W w + Wem p + Wf + Wg (5.9) WING (Ww) + W w = 3.08[((Kwnu,GW)/(t/c)m)((tand, UE - 2(1- A)/A(I+A ))2 + I. 0)XI0_]'sg3(A(I+ A ) )°'_(S)°'741 Notes : K w = 1.00 for fixed wing LE = leading edge sweep angle of the wing GW = Gross Design Weight (ibs) w (t/c)m = maximum thickness ratio nu, = ultimate load factor in g's EMPENNAG_ (Wemp = Wh + Wv + We) (lbs) (5.z7) HORIZONTAL TAIL (Wh) (lbs) W h = .0034 { (WTonu,) 0.813 (Sh) 0.584 (b_tr_h) 0.033 (C/lh) 028)0.915 (5.18) VERTICAL TAIL (Wv) (lbs) W v = 0.19{(I + z_/bv)0"5(WTonua)0"363(Sv)1"0ee(MH) "601 X X (Iv)_'726(i + Sr/Sv)0217(Av)0"337(l + Xv)'363(COS._.c/4v)4)'484} 1"014) N ot ___: w z h = distance from the vertical tail root to where the horizontal tail is mounted on the vertical tail, in ft. Warnina: for fuselage mounted horizontal tails, z h = 0 in ft i v = distance from c/4 to vert. tail Cv/4 in ft.

S r = rudder area ft 2 v = vertical tail taper ratio (5.17) CANARD (We) (ibs) w W c = . 0 0 3 4 ((WTonu,) 0.m3 (S©) 0.¢e4 (bjtr_=) 0.033 (c/i©) 028 )0.915 (s.26) TUSELAGE (W,) (ibs) w_ = 10.4 3 (Ki.,) I_(q_/100)°a_(WTo/Z000)°'_(ig_) 071 Notes : Kin I = 1.25 for airplanes with inlets in or on the fuselage for a buried engine installation.

Kin I = 1.0 for inlets located elsewhere qo = design dive dynamic pressure in Ibs/ft 2 i_ = length of fuselage in ft.

h i = height of fuselage in ft.

(5.4l) _ANDING GEAR GENERAL (Wg) (lbs) Wg = 62.61(WTo/I000) 0"84 (5.42) _ANDING G_AR - _A_N ON W_NG AND NOSE ON FUSELAGE(Wg) Wg = Kg_r(Ag + Bg(WTo) 3/4 + CgWTo + Dg(W70) _2} (ibs) Notes: K0_ r = 1.0 for low wing Kg_r = 1.08 for high wing Constants A through D are in spreadsheet

POWER PLANT (lbs)

Wpw r = W. + W,i + W_ + Wfd + Wp (lbs) ENGINE (W,) (ibs) Actual weight of specific engines Notes: This includes: engine, exhaust, cooling, lubrication.

(6.9) AIR ;NDUCTION SYSTEM (Wai) (ibs) W,, = 0.32 (Nin,) (Ld) (J%n,)°'SS(P2)°s+ (duct support structure) + 1.735{ (Ld) (Ninl) (Ainl)°S(P2) (Kd) (Kin))0.7_I (subsonic part of duct) Notes: K d = 1.33 for ducts with flat cross sections K d = 1.0 for ducts with curved cross sections K m = 1.0 for M D below 1.4 K m = I. 5 for M D above 1.4 L d = duct length in ft.

Nin I = number of inlets Aini = capture area per inlet in ft 2 P2 = maximum static pressure at engine compressor face in psi. Typical values: 15 to 50 psi.

FUEL SYSTEM - SELF SEALING BLADDER CELLS (W_) (ibs) (6.20) W_, = 41.6{ (WF/_,p)/100) 0"818 + W, upp W,upp - 7.91((WF/K_p)/100) 0"_ Notes: _,p = 6.55 ibs/gal for JP-4 (6.26) [UEL DUMPING (Wfd) (ibs) Wfd = 7.38((WF/_,p)I100) °'4s (Ibs) PROPULSION SYSTEM (Wp = W_ + W m + Wwi) (6.23) Engine Controls (W_) (Ibs) fuselage/wing-root mounted jet engines W,c = K,c(i_o) 0_92 Notes: Kec = 0.686 for non-afterburning Kec = 1.080 for afterburning N, = number of engines w if = fuselage length in ft.

b = wing span in ft (6.29) Electric startina System (Wm) (lbs) We, s = 38.93(We/I000) 0"918 Notes: W e = total weight of all engines in Ibs (6.37) water InjectioD (W_) (ibs) Wwi = 8.586W_r/8.35 Notes: W_r = weight of water carried in ibs

FIXED EQUIPMENT WEIGHT (zbs)

Wte q = Wfc + Wel $ + Wiu + W,,_ I + Wox + Wsp u + Wtu r 4. Win.m 4- Wghv 4- + W=_ + W_ + W== Jibs) (7.9) FLIGHT CONTROL SYSTEMS (WI=) (ibs) W1¢ '= _cf (WTo/'10 0 0) 0"581 w 106 for airplanes with elevon control and no horizontal tail 138 for airplanes with a horizontal tail 168 for airplanes with a variable sweep wing (7.19) ELECTRICAL SYSTEMS (W,_) (lbs) W,i , = 426( (W_ + WIN)/1000) °'51 AVIONICS/_NST_UMENTATION/ELECTRONICS (W_) (ibs) Actual Data or Appendix A.

(7.33) AIR/ICE/DE-ICE (W_i) (ibs) W_i = 202( (W_e + 200N=r)/1000) 0"735 (7.39) OXYGEN SYSTEM (W_) (lbs) Wox = 16.9(Nor )1"494 (7.40) APU (W_u) (lbs) W_u = (0.004 to 0.013)WTO (7.47) FURNISHINGS (Wjur) (Ibs) W_r = 22.9(N¢rqD/100) 0"7_ + 107(N=WTo/100,000) °'_ (ejection seats) (misc. emergency equip.)

ARMAMENT (W,,m) (ibs) ACTUAL DATA - APPENDIX A (ibs) GUNS, LAUNCHERS, WEAPONS. (Woh.)

ACTUAL DATA - APPENDIX A (v.so) (ibs) AUXII2tR_Y GEAR (W_=) Waux = O.01W E (v.sl) PAINT ESTIMATE (W_) (ibs) Wpt = 0.0045TO MISCELLANEOUS WEIGHT (W,_) (lbs) W,t c = actual weight data U

AE 622 LIFT STOVL DESIGN

LAST REVISED: Sunday March I0 1990

REVISED BY: Brian Cox

CLASS II COMPONENT WEIGHTS CONSTANTS ..........

..... AIRPLANE GEOMETRY ..... ---WEIGHTS--- 138 L_d 17 31336 K_fcf S_w 322 A_v 1.61 W_TO 1.08 A_inl 3.4 21415 K_ec A__w 3.5 b_v I 0.9 W_E 1P_2(psi) 30 8642 K_g_r LM_LE_w 37.8 S_v 43 W_F I 3557 K_w Lm_w 0.19 l_v 17 W_el 1.25 480 K_inl (t/c)_m_w 0.045 Lm_v 0.35 W_e2 1 W TO.old 225 K_d cbar_w 11.06 LM_c/4_v 38 W_Crew 1.5 31336 21935 K_m b_w 33.57 S_r 0 GW 6.55 0 K_fsp l_f 56 W_wtr W E h_f 6 S_h 40 W_iae 2 21117 b h 11.4 M ff= 0.738 N_inl 630 N_e W_mnz i 420 t_r_h 0.2 W_glw I N_cr W vntv 300 cbar_h 3.75 13.5 0.7 n_ult W_tpipe 300 l_h 16 GW/W_TO 1.2 0.85 M_H W_imas 40 W_mg/W_g W imam 262 W__payload 1196 qbar_D W_E : W_struct + W_pwr + W_feq (2.1) STRUCTURAL WEIGHT W struct = W w + W_emp + W_f + W_g + W__vntv Wing Weight (5.9) W w : 2490 Empennage Weight Vertical Tail (5.18) W v = 256 w Horizontal Tail (5.17) W c = 295 W_emp: Fuselage (5.26 ) W f : 4385 LandinK Gear W_g : 1131 Launch Mechanism W_Imas = W_imam = Ventral Nozzle 300 W_vntv= Therefore, W_struct= 9159 POWER PLANT WEIGHT W__pwr = W_e + W_ai + W_fs + W_fd + W__p + W_mnzl Engine Weizht W e = 4037 Air Induction System (6.9) W ai = 773 I Fuel System Bladder (6.20) W fs = 415 Fuel Dumping (6.26) W fd = 24 Propulsion System W_p = W_ec + W_ess + W_wi Engine Controls (6.23) W_ee = 45.3 Engine Start-Up (6.29) W_ess= 125 Water Injection (6.37) W wi = 0 Therefore, W__p = 170 Main Nozzle Weight W_mnzl = Main Engine Tailpipe section = Therefore, W__pw r : FIXED EQUIPMENT WEIGHT W_feq = W_fc + W_iae + W_els + W_api + W_ox + W_apu + W_fur + W_glw + W_aux + W_pt + W rcsd + W rcsn + W rcsc Flight Control Sys (7.9) W fc = 1021 Avionics (Actual Data) W iae: 1517 Electrical Systems (7.19) W els: 596 Air cond./press./anti- and De-Ice (7.33) W_api= 301 Oxygen System (7.39) W ox = 17 APU (7.40) W_apu: 298 Furnishings (7.47) W fur= 277 Gun and Launch Provisions (Actual Data) W_glw: 630 AUX Gear (7.50) W_aux: 214 m Paint Est. (7.51) W__pt: 204 RCS Duct WeightCWRDC) RCS Nozzle Weight(WRDC) W rcsd= 287 W_rcsn: 83 RCS Controls Weight(WRDC) w m W_rcsc= 35 Therefore, W_feq = 5480 w ............ m : W E = 20777 : m AND BALANCE CLASS II WEIGHT COMPONENT FACTOR WEIGHT x Wx Wz y Wy z 701607 1841719 0 0 160 Fuselaze i 4385 420 0 0 160 398459 1220281 Wing 1 2490 490 60253 Vert Tail i 256 590 151273 0 0 235 = 0 0 160 47190 Hort Tail 1 295 640 188758 w 630919 0 0 138 172410 Main Gear 1.3 1249 505 29764 36378 10 2205 135 Nose Gear 1.3 220 165 Launch Mech 0 0 160 6400 A ASRAAM 40 530 21200 86328 0 0 135 35316 A AMRAAM 261.6 330 41400 173400 0 0 138 Vent Nozzles 1 300 578 0 Z_CE= Struct Wt 1 9498 x_cg: 458 y_cg: 1885210 0 0 165 586905 Engine #I 1 3557 530 0 0 142 68160 110400 Engine #2 1 480 230 173400 0 0 165 49500 #I Tailpipe 1 300 578 277200 0 0 165 69300 #1 Nozzle 1 420 660 293642 0 127502 Air Induct 1 773 380 0 165 161782 0 0 160 66372 Fuel Bladder 1 415 390 = 13230 0 0 140 3368 Fuel Dumping 1 24 550 23799 10 453 150 Eng Controls 1 45 525 0 0 160 19966 68634 Eng Start-Up 1 125 550 0 0 0 0 0 Water Inject 1 0 550 490 163 0 z_cg: Prpl. Wt 1 6139 x_cg: y_cg: Flght Cntrl 1 1021 470 479930 0 0 160 163381 Avionics 1 1517 440 667480 0 0 140 212380 Elect Sys I 596 400 238404 0 0 160 95362 Air/de-ice I 301 460 138250 0 0 160 48087 Oxygen Sys 1 17 155 2620 0 0 150 2535 APU 1 298 600 178615 25 7442 150 44654 Furnishings I 277 125 34594 0 0 160 44280 Gun, Prov. 1 630 330 207900 -35 -22050 140 88200 r AUX Gear 1 214 200 42830 0 0 155 33193 Paint 1 204 420 85547 0 0 160 32589 RCS Duct 1 287 475 136493 0 0 155 44540 RCS Nozzle I 83 475 39545 0 0 155 12904 RCS Controls 1 35 475 16708 0 0 155 5452 Fix Equip Wt I 5480 x_c_: 414 y_cg: -3 z_cg: 151 Empty Weight 21117 x_c_= 456 y_cg= -I z_cg= 157 v Pilot 225 190 42750 0 0 170 38250 W tfo 157 450 70506 0 0 160 25069 ========================================= ............... = ............

Oper Empty Weight 21498 x_c_= 453 y_cg= -I z_cg= 157 Hover Fuel (20%) 1728 410 708655 0 0 160 276548 Fuel 8642 390 3370434 0 0 165 1425953 =========================================== ..........................

W_owe + W_fuel = 30140 x_cZ= 435 y_cg= -0 z_cg= 160 COUNTER AIR MISSION ASRAAM #i 161 530 85330 205 33005 160 25760 ASRAAM ¢2 161 530 85330 -205 -33005 160 25760 AMRAAM #I 327 330 107910 -15 -4905 135 44145 AMRAAM _2 327 330 107910 15 4905 135 44145 AMMO - 400RDS 220 350 77000 0 0 160 35200 BATTLEFIELD AIR INTERDICTION BAI #I 130 72800 Mk-82 #1 560 470 263200 101.5 56840 130 72800 Mk-82 #2 560 470 263200 122.5 68600 130 72800 Mk-82 _3 560 470 263200 -101.5 -56840 130 72800 Mk-82 #4 560 470 263200 -122.5 -68600 130 104910 HARM #1 807 390 314730 62 50034 130 104910 HARM #2 807 390 314730 -62 -50034 22O 350 7700O 0 0 160 35200 AMMO - 400RDS 0 130 24180 EJECTOR RACKS 186 430 79980 0 BAI #2 130 72800 Mk-82 #1 560 470 263200 112 62720 130 72800 Mk-82 #2 560 470 263200 -112 -62720 130 64220 Maverick _1 494 390 192660 62 30628 130 64220 Maverick #2 494 390 192660 77 38038 F 130 6422O Maverick #3 494 390 192660 -62 -30628 130 64220 Maverick #4 494 390 192660 -77 -38038 AMMO - 400RDS 220 350 77000 0 0 160 35200 186 430 79980 0 0 130 24180 EJECTOR RACKS ===================_ ............ . ........... ....------------o CA W_TO 31336 x c_= 433 y_cg= -0 z_cg= 159 BAI _i W_TO 34400 x_cg: 435 y_cg= -0 z_cg: 156 BAI _2 W TO 33642 x_c_= 433 y_cg: -0 z_cg: 157 -0 z_cg: 157 CA W_HOV 24423 x_cg= 447 y_cg= -0 z_cg= 155 BAI #I W_HOV 24382 x_c_= 446 y_cg= -0 z_cg: 155 BAI #2 W_HOV 24744 x_cg= 445 y_cg=

ag2.i ai21..i

The purpose of this appendix is to show the spreadsheets used to calculate the mission performance data and mission capability as discussed in Chapter 8.

Appendix 2: Table of Contents page 2.3 Sustained Load Factor,Turn Rate, and SpecificExcess Energy Calculations .. 301 Brian Cox Last Revised: 12 March 1990 POINT PERFORMANCES: LANDING DISTANCE, SL = 4300 FT TIME TO CLIMB: ALTITUDE 40000 FT TIME 1.75 MINUTES w ABSOLUTE CEILING 80000 FT SPECIFIC EXCESS ENERGY: H a RHO MACH PS (FT) (FT/S) (FT/S) 30000 994.70 0.000889 0.90 505 I0000 1077.40 0.001755 0.90 920 MANEUVERING: = SUSTAINED TURN RATE w H a RHO MACH TURN RATE (FT) (FT/S) (DEG/S) 15000 1057.30 0.001496 0.80 15.00 30000 994.70 0.000889 0.90 10.00 30000 994.70 0.000889 1.20 9.90 OTHER H a RHO MACH LOAD FACTOR (FT) (FT/S) (G'S) 30000 994.70 0.000889 1.60 7.75 15000 1057.30 0.001496 0.90 8.70 ACCELERATION: H _CH TIME START END (FT) (SEC) 30000 0.90 1.60 47.35 30000 0.50 1.40 62.14 I0000 0.30 0.90 18.44 DRAG POLARS: H M CDobase CDo + K*CL^2 CDwave CDmissile CDtrim 0 0.15 0.02277 0.02313 0.11430 0.0000 0.00036 0.00000 0.25 0.02198 0.02234 0.10910 0.0000 0.00036 0.00000 15000 0.56 0.02192 0.02407 0.10190 0.0000 0.00180 0.00035 15000 0.90 0.02230 0.02533 0.10030 0.OOO9 0.00180 0.00033 D 30000 0.85 0.02288 0.02829 0.11030 0.0004 0.00034 0.00462 30000 1.60 0.02034 0.04069 0.10030 0.0196 0.00031 0.00044 0.85 0.02069 0.02173 0.10140 O.00O4 0.00032 0.00027 ;..,J POINT PERFORMANCE WEIGHT 2,82 w 31366 ibs Takeoff Weight = 21498 Ibs Operating Empty Weight = 4321 Ibs 50% fuel weight = 362 Ibs Two short range missiles = 200 rounds ammo = Ii0 Ibs 26291 ibs Performance Weight = 347.90 ft'2 Wing Area = Performance Wing Loading : 75.57 ib/ft^2 Takeoff Wing Loading = 90.16 Ib/ft^2 Takeoff Maximum Thrust = 35573 ib 1.13 Takeoff Thrust to Weight = PERFORMANCE MATCHING CALCULATIONS: #==####_####_##########_###_###_;#######_#_########################_# LANDING DISTANCE: W/S (L) = 5.50*.002378*CL(MAX)*SL W/S (TO) = W/S (L) / {W(LANDING)/W(TAKEOFF)} NOTE: LANDING AT SEALEVEL W(L)/W(TO) = 0.80 154.56 KNOTS V(APPROACH) = CL(MAX) W/S (L) W/S (TO) 1.300 73.11 91.39 #_=_ _#_#._#############_#,,####==######,_############################_ TIME TO CLIMB: RC(SEALEVEL) = HIABS)/T(CL) * LN{ [I- H/H(ABS)]^-I } RC(SEALEVEL) = 528.11 FPS (L/D) (MAX) = .5 * SQRT{ PI*A*e/CDo } (L/D) (MAX) = 12.0 V = SQRT{ [2 * (W/S)] / RHO * SQRT(CDo*PI*A*e) } P(DL) : (L/D)'2 / [ 1 + (L/D)'2 ] P(DL) = 0.99 (T/W) : RC(SEALEVEL)/V / [ P(DL) - SQRT{ P(DL)^2 - P(DL) + [I + (W/S) V (FT/S) (T/W) (MAN) 90.16 395.08 1.36 35635 ibs H = 0.00 ft *********THRUST REQUIRED = M = 0.00 #################################_#####################################= SPECIFIC EXCESS ENERGY: (T/W) (MAN)= PS/V + D/W W = W(TAKEOFF) * W(POINT PERFORMANCE)/W(TAKEOFF) V = MACH * SPEED OF SOUND D = CDo + K*CL^2 CL = [2 * (W/S)] / [RHO * V^2] REQUIREMENT: PS = 505 FT/S MACH = 0.90 H = 30000 FT V = 895.23 FT/S (W/S) CL CD (T/W) (MAN) 75.57 0.2121 0.0298 0.70 H = 30000 ft 18530 ibs *********THRUST REQUIRED = 0.90 M = REQUIREMENT: PS = 920 FT/S MACH = 0.90 H = I0000 FT V = 969.66 FT/S I0000.00 (T/W) (MA (w/s) CL CD 1.21 75.57 0.0916 0.0244 31940 ibs H = 10000 ft *********THRUST REQUIRED = M = 0.90 #_=_#_=#_##_#####################$########_###########################_ MANEUVERING: (T/W) (MAN) = QBAR * _Do / (W/S) + (W/S) * n^2 / [PI*A*e*QBAR QBAR : .5 * RHO * V^2 V = MACH * SPEED OF SOUND n = SQRT{ IV * TURN RATE / G]^2 + 1 } SUSTAINED TURN RATE: H = 15000 FT V = 845.84 FT/S CDo = 0.0253 TURN RATE = 15.00 DEG/S n = 6.95 G'S QBAR = 535.15 PSF CL (w/s) (T/W) (MAN) 0.98 75.57 0.86 22699 ibs H : 15000 ft *********THRUST REQUIRED : 0.8O M : H : 30000 FT 284 V - 895.23 FT/S CDo = 0.0283 TURN RATE - I0.00 DEG/S 4.95 G'S n = 356.36 PSF QBAR = 30000.00 FT CL {W/S> (T/W) (MAN) 75.57 I .05 0.71 18601 ibs H - 30000 ft *********THRUST REQUIRED = M = 0.90 H - 30000 FT V = 1193.64 FT/S 0.0413 CDo = TLrRN RATE = 9.90 DEG/S 6.48 G'S n = 633.53 PSF QBAR = 30000.00 FT CL (W/S) (T/W) (MAN) 0.77 75.57 0.85 S -- 30000 ft *********THRUST REQUIRED = 22320 ibs M = 1.20 H -- 30000 FT V = 1591.52 FT/S CDo = 0.0413 7.75 G'S n = 1126.27 PSF QBAR = 30000.00 FT CL (W/S) (T/W) (MAN) 1.02 0.52 75.57 26810 ibs H = 30000 ft *********THRUST REQUIRED = M = 1.60 H = 15000 FT V = 951.57 FT/S CDo = 0.0253 n = 8.70 G'S QBA_ = 677.12 PSF 15000.00 FT (W/S) (T/W) (MAN) CL 75.57 1.07 0.97 28242 Ibs H = 15000 ft *********THRUST REQUIRED = M = 0.90 _#. # #_######_# -_._-_ _ _ __ .,.. ==,, -#...#.######St##_#..##########_ _####### ########_###-'= ACCELERATION: with Mavg = (Wstart + Wend)/2 ACCEL = (Tstart - Tend)/Mavg TIME = (Vend - Vstart)/ACCEL 1.60 at 30,000 ft 0.90 to REQUIREMENT: M = 895.23 ft/sec Vstart = H = 30000 FT 1591.52 ft/sec Vend = RHO = 0.000889 SLUG/FT^3 V SOUND = 994.70 FT/SEC END CONDITIONS START CONDITIONS W : 26056 Ibs W = 26291 ibs CL : 0.2121 Q : 1126.27 ib/ft^2 CL : 0.0665 Q : 356.36 Ib/ft^2 CD : 0.0411 CD : 0.0332 Thrust : 16085 Ibs Thrust = 4122 lbs FFIow : 61971.60 (ibf/hr) FFIo_ : 5827.00 (ibf/hr) Acceleration = 14.71 ft/sec^2 47.3 sec Time = O. 50 to 1.40 at 30,000 ft REQUIREMENT: M : 497.35 ft/sec Vstart = H = 30000 FT Vend = 1392.58 ft/sec RHO = 0.000889 SLUG/FT^3 V SOUND = 994.70 FT/SEC END CONDITIONS START CONDITIONS W : 26016 ibs W = 26291 Ibs Q : 862.30 ib/ft_2 CL : 0.6871 CL : 0.0867 Q : 109.99 ib/ft^2 CD : 0.0414 CD : 0.0804 Thrust : 14785 ibs Thrust = 3075 lbs FFlow = 55390.00 (ibf/hr) FFlow = 2861.00 (Ibf/hr) 14.41 ft/sec^2 Acceleration : 62.1 sec Time : 0.30 to 0.90 at I0,000 ft REQUIREMENT: M : H : I0000 FT Vstart : 323.22 ft/sec RHO : 0.001755 SLUG/FT^3 Vend : 969.66 ft/sec V SOUND : 1077.40 FT/SEC L_ END CONDITIONS START CONDITIONS W = 26101 ibs W : 26291 ibs CL : 0.8243 Q = 825.06 Ib/ftA2 W CL : 0.0909 Q : 91.67 Ib/ftA2 CD : 0.1032 CD : 0.0415 Thrust = 31834 ibs Thrust = 3293 lbs w FFlow = 73249.00 (ibf/hr) FFlow = 2861.00 (Ibf/hr} Acceleration : 35.05 ft/sec^2 18.4 sec Time : SUMMARY OF POINT PERFORMANCE THRUST H M W MAN REQUIREMENTS (T/W) MAN 1.14 31366 IA 0.00 0.00 40000.00 1.75 min 80000.00 0.70 26291 2A 30000 0.90 505 ft/sec 1.21 26291 2B I0000 0.90 920 ft/sec 0.86 26291 3A 15000 0.80 15.00 deg/sec 0.71 26291 3B 30000 0.90 10.00 deg/sec 0.85 26291 3C 30000 1.20 9.90 deg/sec 1.02 26291 3D 30000 1.60 7.75 g's 1.07 26291 3E 15000 0.90 8.70 g's 47.3 sec 0.61 26291 4A 30000 0.90 1.60 62.1 sec 0.56 26291 4B 30000 0.50 1.40 18.4 sec 1.21 26291 4C 10000 0.30 0.90 w MISSION ANALYSIS FOR AE 622 Brian Cox Last Revised: I0 April 1990 The following mission legs burn fuel and need to be accounted for: Battlefield Air Int.

Counter Air Superiority I. Engine Start/Warm Up I. Engine Start/Warm Up 2 Taxi 2. Taxi 3 Short Take-off 3. Short Take-off 4 Accelerate to Climb Speed 4. Accel. to Climb Speed 5 Cllmb 5. Climb 6 Subsonic Cruise 6. Subsonic Cruise 7. Sea-Level Dash-in 7 Accelerate to Supersonic Cruise 8. Strafe Run 8 Supersonic Cruise 9 Combat 9. Sea-level Dash Out i0. Supersonic Cruise I0. Climb II. Subsonic Cruise Ii. Subsonic Cruise 12. Hover - Half Minute 12. Hover - Half Minute 13. Landing 13. Landing For the following segments, statistical fuel burns fractions are used due to lack of detailed analsis methods (at this time): * Engine Start/Warm-up * Taxi * Short Takeoff t Landin_ Also, the followinK is assumed as part of the mission: CA mission combat fuel burn: 20 % of total fuel BAI mission strafe run fuel burn: 10 % of total fuel Equations and methods specific to the aircraft flight phase under consideration will be used for the remaining mission legs.

Aircraft Parameters CA Mission Takeoff Weight = 31366 Ibs BAI Mission Takeoff Weight = 34400 Ibs 8642 Ibs Fuel Weight = 347.9 ft^2 Wing Area, S = CA Mission, WOE = 21498 ibs BAI Mission, WOE = 21778 Ibs 20% fuel = 1728.4 Ibs 10% fuel = 864.2 Ibs 1196 ibs 2328 ibs Hover Weapons = Hover Weapons = 24422 ibs Hover Weight = Hover Weight = 24970 Ibs 1.27 1.27 (T/W) Required = (T/W) Required = Hover Thrust = 31016 ibs Hover Thrust = 31712 ibs Drag Polars BAI #I CA Mission CDo ÷ K*CLA2 CDo + K*CL^2 H (ft) Mach 0.02277 0.1143 0 0 15 0.02277 0.I143 0.02198 0.1091 0 0 25 0.02198 0.1091 15000 0 56 0.02372 0.1019 0.02372 0.1019 15000 0 90 0.02500 0.i003 not in mission 30000 0 85 0.02790 0.1103 0.02790 0.1103 not in mission 30000 1 60 0.04038 0.1003 0.02136 0.1014 0 0 85 0.02136 0.1014 COUNTER AIR MISSION I. Engine Start/Warm Up ************************** Fuel Burn : 314 ibs WI/WTO = 0.99 ************************** Wl : 31052 ibs WF = 8328 ibs 2. Taxi ************************** Fuel Burn = 279 ibs W2/WTO = 0.99 ************************** W2 : 30773 ibs WF : 8049 Ibs 3. Short Takeoff Using 0.5 minutes for takeoff thrust setting: 0.008333 hrs LIFT Engine: CRUISE Engine: SFC : 0.8086 (Ibf/hr)/ibt SFC : 1.311 (ibf/hr)/ibt T(TO) : 12500 ibs T(TO) = 25250 Ibs WFDOT = 10107.5 Ibf/hr WFDOT =33102.75 Ibf/hr FuelBurn= 84 ibs FuelBurn= 276 ibs Fuel Burn = 360 ibs W3 = 30413 ibs WF : 7689 Ibs 4. Accelerate to Climb Speed (Out) Accelerate from M= 0.20 to M = 0.80 at sea level, so Vstart = 533 ib/ftA2 223 ft/sec q(3/4 V)= Vend = 893 ft/sec Acceleration = 25 ft/sec^2 26.8 sec t(acc) = Thrust Required = Acceleration Force + Drag Using H=0, M=0.25 drag polar, the begin weight, and velocity at 75% Vend throughout the acceleration: Drag increment for two short range missiles: 0.00018 CL = 0.1639 CD = 0.0251 236311bs Acceleration Force, F = Dra_, D = 4657 ibs 28288 ibs Thrust Required, T = From Engine Deck, SFC = 1.485 (ibf/hr)/ibt Fuel Flow, WFDOT = 42008 Ibf/hr Fuel Burn = 313 Ibs W4 : 30100 ibs WF : 7376 ibs 5. Climb (Out) Average Rate of Climb = 15000 ft/min Time to Climb to 30000 ft = 0.033 hrs Use climb variables at 2/3 final altitude: M = 0.80 V = 829 ft/sec qbar = 436 Ib/ft^2 The aircraft travels horizontally 99533 ft while vertically 30,000 ft so theta is 16.77 degrees.

So, L = 28820 Ibs D = 8686 Ibs CL = 0.1902 Usin_ dra_ polar for H=15000ft, M=0.56 for H=20000ft, M = 0.8 Drag increment for two short range missiles: 0.000174 CD = 0.0276 D = 4180 ibs 29O Thrust Required, T = 12866 ibs From Engine Deck, SFC = 1.132 (Ibf/hr)/ibt Fuel Flow, WFDOT = 14564 ibf/hr Fuel Burn = 485 Ibs W5 = 29615 ibs WF = 6891 Ibs 6. Subsonic Cruise (Out) Ranze = I00 - 16 = 84 nm (climb range credit) Cruise Mach Number : 0.80 qbar : 282 Ib/ft^2 CL = 0.3023 Drag increment for two short range missiles: 0.000168 CD = 0.0381 3737 Ibs Dra_ = Thrust Required, T = From Engine Deck, SFC : 0.801 (ibf/hr)/ibt Fuel Flow, WFDOT = 2993 Ibf/hr Cruise Time = 0.177 hrs Fuel Burn = 531 ibs W6 = 29084 ibs WF = 6360 ibs 7. Accelerate to Supersonic Cruise (Out) Accelerate from M= 0.80 to M = 1.60 at sea level, so Vstart = 774 ft/sec q(3/4 V)= 600 Ib/ft^2 Vend = 1549 ft/sec Acceleration = 17 ft/sec^2 45.6 sec t(acc) = Thrust Required = Acceleration Force + Drag Using H=30 k ft, M=1.6 drag polar, the begin weight, and velocity at 75% Vend throughout the acceleration: Drag increment for two short range missiles: 0.000155 CL = 0.1393 CD = 0.0425 15367 ibs Acceleration Force, F = 8869 ibs Dra_, D = r 24236 ibs Thrust Required, T = 2.023 (ibf/hr)/ibt From Engine Deck, SFC = 49030 ibf/hr Fuel Flow, WFDOT = Fuel Burn = 620 Ibs W4 = 28463 Ibs WF = 5739 ibs 8. Supersonic Cruise (Out) Range = 50 nm Cruise Mach Number : 1.60 qbar : 1126 Ib/ft^2 CL = 0.0726 Drag increment for two short range missiles: 0.000155 CD : 0.0411 w 16090 Ibs Dra_ : Thrust Required, T : = 1.563 (Ibf/hr)/ibt w From Engine Deck, SFC : 25149 ibf/hr Fuel Flow, WFDOT : Cruise Time : 0.053 hrs Fuel Burn : 1334 ibs ************************** W6 : 27130 ibs 4406 ibs WF : 9. Combat Usin_ 20% total fuel for combat: Fuel Burn = 1728 ibs W9 : 25401 ibs WF : 2677 Ibs Dropping two ASRAAMS and half ammo: W9 : 24969 ibs

WF = 2677 Ibs

I0. Supersonic Cruise (In)

Ranze = 50 nm

Cruise Mach Number = 1.60 1126 ib/ft^2 qbar = CL = 0.0637 CD = 0.0408 Drag = Thrust Required, T = 15982 ibs From Engine Deck, SFC = 1.563 (ibf/hr)/ibt Fuel Flow, WFDOT = 24979 Ibf/hr Cruise Time = 0.053 hrs Fuel Burn : 1325 Ibs Wll = 23645 ibs WF = 1353 ibs II. Subsonic Cruise {In) Ranze = I00 nm Crulse Mach Number = 0.80 qbar = 282 lb/ft^2 CL = 0.2414 CD = 0.0343 3363 lbs Drag = Thrust Required, T = From Enzine Deck, SFC : 0.801 (Ibf/hr)/ibt Fuel Flow, WFDOT = 2693 Ibf/hr Cruise Time = 0.212 hrs = w Fuel Burn = 571 lbs W12 = 23074 Ibs WF = 782 Ibs 12. Hover Half Minute : 0.0083 hrs Hover Thrust: 293 12105 ibs LIFT Engine = 18911 ibs MAIN Engine = From Enzine Deck, SFC: LIFT Engine = 0.8094 (ibf/hr)/Ibt MAIN Engine = 0.924 (ibf/hr)/Ibt Fuel Flow, WFDOT: LIFT Engine = 9798 lb/hr MAIN Engine = 17474 lb/hr Fuel Burn, Wburn: LIFT Engine = 82 Ib MAIN Enzine = 146 Ib _z Fuel Burn = 227 Ibs WI3 = 22846 ibs WF = 554 ibs 13. Landing $$$1Z*$*1111111*1$$15*111I 0.995 W14/W13 = Fuel Burn = 114 lbs 22732 lbs W14 = WF = 440 lbs CA Mission Fuel Burn Summary-- Fuel Burn Phase 314 lbs I. Engine Start/Warm Up 2. Taxi 279 lbs 3. Short Take-off 360 lbs 313 lbs 4. Accel. to Climb Speed 5. Climb 485 lbs 6. Subsonic Cruise 531 lbs 620 lbs 7. Accel. to Supersonic Cruise 1334 lbs 8. Supersonic Cruise 9. Combat 1728 lbs 1325 lbs 10. Supersonic Cruise II. Subsonic Cruise 571 lbs 12. Hover - Half Minute 227 lbs 13. Landin_ 114 lbs CA Mission Fuel : 8202 ibs -BATTLEFIELD AIR INTERDICTION MISSION I. Engine Start/Warm Up =$***$$t$_$,tt,$,tt*$tzsz* Fuel Burn = 327 ibs Wl/WTO = 0.99 Wl = 34073 ibs WF = 8315 ibs 2. Taxi Fuel Burn = 307 ibs W2/WTO = 0.99 W2 : 33767 ibs WF : 8009 Ibs 3. Short Takeoff 0.5 minutes for takeoff thrust setting: 0.008333 hrs Usin_ CRUISE Engine: LIFT Engine: SFC = 1.311 (ibf/hr)/Ibt SFC : 0.773 (Ibf/hr)/ibt T(TO) : 12500 ibs T(TO) = 27000 Ibs WFDOT = 9668.75 ibf/hr WFDOT = 35397 ibf/hr FuelBurn= 295 ibs FuelBurn= 81 ibs Fuel Burn : 376 ibs W3 : 33391 ibs WF : 7633 Ibs 4. Accelerate to Climb Speed (Out) Accelerate from M: 0.20 to M : 0.80 at sea level, so 533 ib/ft^2 Vstart : 223 ft/sec q(3/4 V): Vend = 893 ft/sec Acceleration = 24 ft/sec^2 27.9 sec t(acc) : Thrust Required : Acceleration Force + Drag Usin_ H=0, M=0.25 drag polar, the begin weight, and velocity at 75% Vend throughout the acceleration: 0.00331 Drag increment for BAI #i mission: CL = 0.1799 CD = 0.0288 Acceleration Force, F = 24908 Ibs 5349 ibs Draz, D = 30257 Ibs Thrust Required, T = From Engine Deck, SFC = 1.311 (ibf/hr)/Ibt Fuel Flow, WFDOT = 39667 Ibf/hr Fuel Burn = 308 ibs ************************** W4 : 33083 ibs WF : 7325 ibs 5. Climb (Out) Average Rate of Climb = 15000 ft/min Time to Climb to 30000 ft = 0.033 hrs Use climb variables at 2/3 final altitude: M = 0.80 V = 829 ft/sec qbar : 436 ib/ft^2 The aircraft travels horizontally 99533 ft 16.77 degrees.

while vertically 30,000 ft so theta is So, L : 31676 Ibs D : 9547 ibs CL : 0.2090 Usinz draz polar for H:15000ft, M:0.56 for H:20000ft, M : 0.8 Drag increment for BAI #I mission: 0.00339 CD : 0.0316 u D : 4783 Ibs 143311bs Thrust Required, T : 1.126 (ibf/hr)/Ibt From Engine Deck, SFC : 16136 Ibf/hr Fuel Flow, WFDOT : ************************** Fuel Burn = 538 Ibs ************************** W5 : 32546 ibs WF : 6788 ibs 6. Subsonic Cruise (Out) 184 nm 16 = Range = 200 (climb range credit) 0.80 Cruise Mach Number = 282 ib/ftA2 qbar = CL = 0.3322 0.0032 Drag increment for BAI #I mission: 0.0433 CD = 4239 ibs Dra_ = Thrust Required, T = 0.806 (Ibf/hr)/ibt From Engine Deck, SFC = 3417 ibf/hr Fuel Flow, WFDOT = 0.389 hrs Cruise Time = Fuel Burn : 1331 ibs W6 = 31215 ibs WF = 5457 ibs 7. Sea Level Dash (Out) Range = 80 nm m Cruise Mach Number = 0.85 10711b/ft^2 qbar = 0.0838 CL = 0.00309 Drag increment for BAI #I mission: CD = 0.0252 9373 ibs Dra_ = Thrust Required, T = From Engine Deck, SFC = 0.903 (Ibf/hr)/Ibt w Fuel Flow, WFDOT = 8463 ibf/hr Cruise Time = 0.142 hrs Fuel Burn = 1204 Ibs ************************** W7 = 30011 ibs WF = 4253 ibs 8. Strafe Run Usin_ 10% total fuel for strafe run: ************************** Fuel Burn = 864 ibs W8 = 29146 Ibs WF = 3388 Ibs Dropping two Mark 82 Bombs and two AGM 65's: W8 = 27038 Ibs WF : 3388 Ibs 9. Sea Level Dash (In) 80 nm Range : m 0.85 Cruise Mach Number : 10711b/ft^2 qbar = 0.0726 CL : 0.0013 Drag increment for two AGM 65: CD = 0.0232 8640 Ibs Drag : Thrust Required, T : 0.903 (ibf/hr)/ibt From Engine Deck, SFC : Fuel Flow, WFDOT : 7802 ibf/hr w Cruise Time : 0.142 hrs i Fuel Burn : III0 Ibs m W9 = 25928 ibs WF : 2278 ibs I0. Climb (In) Average Rate of Climb : 15000 ft/min 0.033 hrs Time to Climb to 30000 ft : Use climb variables at 2/3 final altitude: M : 0.80 V : 829 ft/sec qbar : 436 Iblft^2 The aircraft travels horizontally 99533 ft while vertically 30,000 ft so theta is 16.77 degrees.

L : 24825 ibs So, D : 7483 Ibs CL : 0.1638 Usin_ drag polar for H:lS000ft, M:0.56 for H:20000ft, M : 0.8 w 0.0014 Drag increment for two AGM 65: CD = 0.0279 D = 4221 ibs Thrust Required, T = 11704 Ibs w From Engine Deck, SFC = 0.836 (ibf/hr)/ibt Fuel Flow, WFDOT = 9784 ibf/hr Fuel Burn = 326 ibs WlO : 25602 Ibs WF : 1952 ibs II. Subsonic Cruise (In) Range = 200 16 = 184 nm (climb range credit) Cruise Mach Number = 0.80 qbar = 282 lb/ft^2 CL = 0.2614 Drag increment for two AGM 65: 0.00134 CD = 0.0368 3602 ibs Drag = Thrust Required, T = From Engine Deck, SFC = 0.801 (ibf/hr)/ibt Fuel Flow, WFDOT = 2885 ibf/hr Cruise Time = 0.389 hrs Fuel Burn : 1124 ibs Wll = 24479 ibs u WF = 829 ibs w 12. Hover Half Minute = 0.0083 hrs Hover Thrust: LIFT Engine = 12800 Ibs MAIN Engine = 18912 ibs From Engine Deck, SFC: LIFT Engine : 0.8086 (Ibf/hr)/ibt MAIN Engine = 1.01498 (ibf/hr)/ibt Fuel Flow, WFDOT: 10350 ib/hr LIFT Engine = }JAIN Engine = 19195 ib/hr Fuel Burn, Wburn: 86 Ib LIFT Engine = 160 ib MAIN Engine = Fuel Burn = 246 ibs W13 = 24232 ibs WF = 582 Ibs 13. Landing ************************** Fuel Burn = 121 ibs 0.995 WI3/WI2 = ************************** 241111bs W13 = 461 Ibs WF = BAI Mission Fuel Burn Summary ....

w Phase Fuel Burn 327 Ibs I. Engine Start/Warm Up 307 Ibs 2. Taxi 376 ibs 3. Short Take-off 308 Ibs 4. Accel. to Climb Speed 5. Climb 538 Ibs !W 6. Subsonic Cruise 1331 ibs 1204 ibs 7. Sea-Level Dash-in 8. Strafe Run 864 Ibs n 9. Sea-level Dash Out IIi0 Ibs 326 Ibs I0. Climb 1124 ibs II. Subsonic Cruise 246 Ibs 12. Hover - Half Minute 121 ibs 13. Landing BAI Mission Fuel = 8181 ibs w 3OO

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--= ORIGINAL PAGE iS OF POOR QUALI'I7f ?KRFO_ _R T_ _OH_CH - D(_ _O_EPLOT 14 April 1990 Bri_ Coz W : 20291.00 (i/8J : S : 347.90 a = 9 : 15000 DELTA : 0,5643 _C:6 _o ! _kR CL ntu,,=: f _TE 0A64 0.0241 9.053 58 1.3001 0.2108 Z2305,60 Z,4C 1.30 1.00 I,)0 0.00 0.3 0.0238 9.243 75 1.0045 0,1330 22811.40 2.79 1.25 1,25 1.25 4,35 0.4 0.0235 9.434 134 0,5650 0.0574 23317.20 3,70 I.Z1 2,14 ;.14 d,25 0,5 0.0233 9.024 209 0,36i60.0368 24205.90 4.?7 1,17 3.22 3.22 10,88 4.46 (.¢6 12.83 0._ 0.0229 9.814 301 0,25ii0,0293 25094,¢0 5,81 1.12 5.83 5,s3 14,30 0.7 0.0226 9.892 410 0,18450.0281 2621?,00 6,85 1,08 0,8 0,02Z3 9.970 535 0.14130.0243 2?339.80 7.89 1,03 ?,29 _._9 15.74 0.9 0.0232 9,970 6?7 0.1116 0.0244Z8550.40 8.86 0,98 8.78 6,?) 16.90 1.0 0,0373 9.9?0 836 0,09040,0381 29529,60 6,85 0.93 10.29 !.65 15.34 i.i C.0395 }0.136 1011 0,01470,040i 30556,80 9,27 0,89 11.84 ).J0 14,18 _1_ _,)418 10.301 1204 0.06Z80.042_ 31584,00 9,38 0.84 13,38 :.)0 13.00 1.3 O.J4iO 10,467 1413 0.0535 0.041_ 32860.2U 9.73 0,73 13.+4 9.)U 12,00 !,4 ),el01i0.633 1638 0.0461 0,0403 34136.40 9.93 0.63 13,66 _,00 11,14 i.5 0.0398 10,799 1881 0.0402 0.0399 35502.00 9.84 0.59 14.6_ LDO 10.40 t,6 0,039410,964 2140 0.0353 0,0395 36867,60 9,12 0.55 15,5T LO0 9,75 3O5 ORIG,+H_L PAGE IS OF POOR QUALITY' MISSION PEIIFO_NCE DR AE 6ZZ- FERRY liANOH CALCULATIONS _rmn gox La._ hvised: 14April 1990 ............ Aircraft Farameters CA Mission T_Weoff Weight : 33076 Ibs BAIMission TueoffWeight : 3440U Ibs PuelWeiiht : 1035_ Ibs Win{Area, H : 347.b ft"Z CAMission, _E : Z1498 ibs 2U%fuel: Z07_.4 IDs Hover Weapons : I19_!bs hover Weight = 24764 Ibs :Iik: Required = I,Z7 Hover hrust = 314511_ [,ra_ Polars CA Mission H Hi> Math _c + R:CL'Z 0,02377 0,1143 u 0.15 g.OZ27? 0.1143 0,02198 0,1091 U 0.2_ 0.02198 0,1Obl 0.0237Z 0,1019 150o0 0.56 0,02372 0,1019 15o0_ 0,9u 0.0250u 0.I003 _otinmi_slon 30000 0.85 0,0279u 0,1103 0,02790 0,1103 30U0_ 1.60 0.04038 0.100_ notinmission 0,0Z136 0,1014 0 0,85 0,02136 0,1014 ===========BEST MACHAN£' ALTI_E STUDY w llllX{{{}{{{lllll{$}{{{}ll 0.99 Fuel Burn: 198|bs llll{$lllllllllllllil{{lll WI : 32878 Ibs WF : 10154 Ibs 2,TAXi l}}{llllli{$}l}{llll}}}}}!

FuelBurn: 184Ibs W2/W'PO : 0,99 ll{IIIIUIII{{{{lUIII{{H W2 : 32713 Ibs WF: 9989 Ibs 3,TAE_3F_ 34)6

Of POOR QUALITY

Osin£ u.35 minutes for t_keoff thru_ settle: 0,005833 hrs LI__ _ine: CRUI_E Rnline: SFC: 0.80_6 [lbf/br)/lbt S_ = 1.311 Ilbf/hr)/lbt T(TO) : 12500 Ibs T(TO_ : 25250 lbs WF_T : 10107.5 Ibf/hr k{P_T :3310Z.Y5 Ibflhr FuelBurn: 5_ Ibs FuelBmm: 193Ibs ui{$$l$}{{${{${{${{{{{{{!

FuelBum : 252 Ibs mmmmmmmm,, WZ : 324611bs WF : 9737Ibs 4.ACCEL_Ti T_CL!M_ Accelerate fromM: 0.20 to M : 0.80 at sealevel, so Vstart : 223 ft/sec q(_t4 VI: 533 lb/ft'i Vend: _93 ft/sec Acceleration : 25 ft/sec'Z w_ tiacol : Z8.8 sec ,nrus. Require_ : Acceleration _orce • {)r_ w Us_n_ H=0,M:_.Z5 dr_ _l_, the beIin weight, and vel_ity at ?5%Vend throughout theaooeleratlon: Dra_increment fortwo short ruge mis{iles: 0.00018 CL : 0.i_4_ _ : 0.0255 Acceleration Force, F : fiZZ3 Ibs Or_, _ : 4?32Ibs Thrust F_quired, T : 29955 Ibs = = m From h_ine 1)_k, SFC: 1,485 (Ibflhrlllb_ FuelFlow, _OT : 44484 Ibflbr llltllll$lllllll$ll$$tllll FuelBurn: 331Ib{ w {I$$III$III$llIIIUIIII$ll W4: 32130 lbs WF: 9406 lbs 5. OLIM_ m Average HateofOlimb= 25000 ft/min TimetoClimb to30000 ft = 0,020 hrs 30"/ Useclimb variables at 2/3final mltitude: ORIGINAL PAGE _S OF POOR QUALITY M : 0.8_ V : 82_ fttsec qbar = 436 ]b/ft'2 59?2Oft Theaircraft travels borizontally _iie vertically 30,000 ft so theta is 26.6? degrees, L : 28711 lbs SO, D : 14423 lbs CL: 0,1894 _ir_ dra_ polar for _:lS000f_, M=0.56 for E=20000ft, M : 0.8 Dra_ increment for two short range missiles: 0.000174 Ob: 0,027_ D : 4175 lbs 1859_ lbs 1,132 Ilbf/hr)/lbt Fuel Flow, g_30T: 21053 ]bf/hr $ttJZttXZ{S{_tZt$1$11$$_t{ Fuel Burr : 421 Ibs IIIllIIIIUlIIIIIIIIiIIIII W5 : 31709 lb_ WF = 8985 ibs 6, S_SON!C _2UISB I.E. FE_Y RANGE FERRY RAN_ WINULOADING : 80,20 m H : 30000 _ a : 995 MACH _BAR CL CDo PISAZe TR_ FUELFLOW TIN{ _ (_ 0,40 ?0 1,139 0,0241 8.184 4209 2976 10730 ?03 0,50 Ii0 0.129 0,0238 8,854 3208 2610 12232 lO01 0,60 158 0.506 0,0235 8,923 2878 2245 14224 1397 0,70 216 0,372 0.0232 8.995 Z894 Z510 12720 1458 0,80 Z8Z 0.285 0,0229 9,086 3120 2776 11505 1507 0,90 356 0,225 0.0238 9.086 3843 3506 9108 1342 1,00 440 0.182 0.0379 9.086 8382 8440 4958 812 1.I0 532 0,151 0.0435 9.217 8512 9254 3450 621 1,20 834 0,127 0,04Z3 9,367 9700 12069 164_ 520 1,30 744 0,108 0.0414 9,518 11025 15553 _053 437 1,4_ 862 0,093 0,040? 9,869 12478 19037 1677 384 1.50 990 0,081 0.0404 9,819 14143 _250 1435 352 1,80 1126 0,071 0.0399 9,9?0 15833 25462 1254 328 3O8 ORIG_AL P,_G_" iS OF POOR QUALITY H : 36089FT a : 968 PI:k:e FURL FLOW TI_ MACH QBAR CL CDo 0,50 83 0.989 0.0241 4,417 6817 2723 11728 934 0.60 119 0.673 0.0238 8.834 3112 2368 13487 1Z89 0,70 ]62 0,495 0.0234 8.905 Z870 2430 13140 1486 0,80 Z12 0,379 0,0231 8,975 2879 2493 12811 1633 0.90 288 0.299 0.0240 8.9Y5 3168 2928 10908 1584 1.00 33t 0,242 0.0381 8.975 5140 5095 8268 999 ].lO 400 0,200 0.0437 0.125 6700 1084 4508 790 I.ZO 477 0.168 0,0426 9.274 7569 9073 3519 873 1,30 559 0,143 0.0416 9.423 8519 11529 2770 574 1,40 849 0.124 0,0409 9.572 9590 13904 2283 509 ].50 745 0,108 0.0406 9.721 10826 16283 1963 469 1,60 847 0.095 0,0401 9,870 12086 18541 1722 439 H : 40000 Ff a : 988 MACH QBAR CL ¢Do PI:A$e VOEL FLOW TI_ 0.5_ 69 1,170 0.0243 4.3?3 8044 3008 10816 846 0,80 99 0,812 0,0240 8,745 3416 2818 12208 1167 0.70 134 0,59? 0,0237 8.816 2995 2554 12503 1395 0,80 175 0.457 0.0233 8.886 2860 2492 12813 1633 0.90 222 0,361 0.0242 8,886 3007 2794 11429 1839 1,00 274 0.292 0,0383 8.886 4576 4559 7004 1118 1,10 33Z 0.242 0.0439 9,033 5815 8176 5170 906 1,20 395 0.203 0,0428 9.181 8493 7794 4097 783 1,30 463 0.173 0,0418 9.329 7263 9838 3248 672 1,40 537 0,149 0.0411 9,476 8126 11881 2688 599 1,50 817 0,130 0,0408 9.624 913U 13773 2319 554 1.60 702 0,114 0.0403 9.772 10179 15662 2039 520 R : 45000 I;'I' a : 988.1 PItAle MAC{] QSAIi CL Ol)o 0,60 78 1.034 0,0242 8.658 3985 4615 8919 $61 0.70 106 0,759 0,0239 8,727 3306 3814 8838 986 0.8_ 138 0.581 0.0236 8.797 2975 2812 12223 1558 0.90 175 0,459 0,0245 8.797 Z943 2756 11587 1862 1.00 216 0.372 0,0386 8.T97 4073 4075 {836 I449 1.I0 261 0,308 0,0441 8,943 4964 5307 6017 1055 1.20 310 0.258 0.0430 9.089 5438 8540 488_ 934 1.30 364 0,220 0.0421 9,235 5996 8139 3923 813 1,40 422 0.190 0.0413 9.381 6638 9738 32?9 731 3O9 OR{GINAL PAGE IS OF POOR QUALITY 1.50 485 0.165 0.0410 9.528 7400 11203 2950 661 1.6_, 552 0.145 0.0406 9.674 8204 12669 2621 643 968,1 _o PIZASe _ P_TELFI,O_ TI_ _ I_!

0.80 109 0,739 0.0238 8.709 3286 3137 10179 1297 0,90 137 0,584 0,0247 8,709 3051 3828 11291 1619 1.00 170 0,473 0,0388 8.709 3805 4080 ?826 1247 1.10 205 0,39] 0.0444 8,853 w 4399 5336 5985 1049 1.20 244 0,328 0.0432 8,998 4691 8591 4845 92_ 1,30 287 0.280 0.0423 9,143 5071 7370 4333 89?

1,4D 332 0.241 0,0415 9.288 5530 6150 3918 874 1.50 382 0.210 0,0412 9.432 6092 9265 3446 824 ],60 434 0.185 0,0408 9.577 6696 I03_i 3078 784 OR1GIN._L PAGE IS OF POOR OUALITY MISSION CAPABILITY ANALYSIS 19 April 1990 Brian Cox 8642 Ibs Total Mission Fuel 953 ibs Takeoff 313 ibs Acclerate to Climb at Sea Level 627 ibs Accelerate to Supersonic Cruise w 227 ibs Hover 114 Ibs Landing 432 Ibs Reserves Climb {ib fuel/ft) 0.0162 Ibs/ft Subsonic Cruise (Ib fuel/nuatical mile) 5.711bs/nm Supersonic Cruise (ib fuel/nautical mile) 26.70 ibs/nm Low Level Dash (Ib fuel/nautical mile) 15.05 ibs/nm w MASS INTERCEPT Fuel Burn 953 ibs Takeoff 880 ibs Climb to 35000 ft 627 Ibs Accelerate to Supersonic Cruise 2670 ibs Supersonic Cruise for 115 nm Dash at 5000 ft for 30 nm 452 ibs Combat with K = 0.15 % of total fuel 1296 Ibs Climb to 30000 ft 486 ibs 485 ibs Subsonic Cruise for 85 nm 773 ibs 9. Hover, Landing, and Reserves Total Fuel Burn 8622 ibs HIGH VALUE ASSEST PROTECTION Fuel Burn 953 Ibs 1 Takeoff 1042 ibs 2 Climb to 45000 ft 5 Loiter for 0.506553 hrs 1042 Ibs 3 627 Ibs Accelerate to Supersonic Cruise 4 1335 ibs Supersonic Dash for 50 nm 5 1107 ibs Shoot Missiles at Optimum Climbing Turn L 6 1335 ibs Supersonic Dash for 50 nm 7 Subsonic Cruise for 75 nm 428 ibs 773 ibs 8. Hover, Landing, and Reserves 7600 Ibs 8642 ibs Total Fuel Burn TRANSPORT {HELICOPTER) INTERCEPT Fuel Burn 953 Ibs 1. Takeoff 799 ibs 2. Climb to 30000 ft 627 Ibs 3. Accelerate to Supersonic Cruise 4406 ibs 4. Supersonic Cruise for 180 nm 6. Combat with K = 0.15 % of total fuel 1296 ibs 7. Climb to 30000 ft 486 Ibs 8. Subsonic Cruise for 180 nm 1028 ibs 773 Ibs 9. Hover, Landing, and Reserves Total Fuel Burn 10368 ibs w ..................... m-- STOVL TWO STAGE MISSION Fuel Burn PHASE 1 953 ibs I. Takeoff 799 ibs 2. Climb to 30000 ft 1370 Ibs 240 nm 3. Subsonic Cruise for 773 ibs 4. Landing, Hover, and Reserves PHASE 1 Fuel Burn 3895 Ibs PHASE 2 953 Ibs 5. Takeoff 799 ibs 6. Climb to 30000 ft 1068 Ibs 7. Supersonic Dash for 40 nm __= 0.15 % W Fuel 1296 Ibs 8. Sea Level Combat/Strafe Run at K= 9. Climb to 30000 ft 799 ibs 1602 ibs 9.5 Supersonic Cruise for 60 nm 10. Subsonic Cruise for 240 nm 1370 ibs 773 ibs II. Hover, Landing, and Reserves 8661 Ibs PHASE 2 Fuel Burn

APPENDIX 3 The purpose of this appendix is to present the stability and control engineering calculations for the material presented in Chapter 9.

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Fllght Condition 2 M =0.85 h= 100 ft.

Open Loop 0.050 Short Period Frequency radls_ ii 0.160 Short Period DampingRatio Dutch Roll Frequency 0.560 radlsec 0.180 Dutch Roll Damping Ratio OmegaS,P,/Roll Rate OmegaD/Roll Rate

OmegaD/Omega S, P. p l (c_Isec)

0.01 1,0290909090909 0.II 286,478220 0.05 1,0303 i 57894737 0.56 57.2956455 0,I0 1.032 1,12 28.6478228 0,15 1.0338823529412 1,68 19.0985485 0,20 1.036 2.24 14,3239114 0,25 1.0384 2,80 11.4591291 0.30 1.0411428571429 3.36 9.54927426 0.35 1.0443076923077 3,92 8.18509222 0.40 1.048 4.48 7,16195569 0.45 1.0523636363636 5.04 6,36618284 0.50 1,0576 5.60 5,72956455 0.55 1.064 6.16 5,20869505 0.60 1.072 6.72 4,77463713 0.65 1.0822857142857 7.28 4,40735735 0,70 1,096 7.84 4,09254611 O,75 1.1152 8.40 3.8197097 0.80 1,144 8.96 3.58097785 0.85 1.192 9.52 3.37033209 0.90 1.288 0.08 3,18309142 0.91 1.32 0.19 3.14811239 0.92 1.36 0,30 3,11389378 0.93 1,4114285714286 0.42 3,08041105 094 1.48 0.53 3,04764072 0.95 i.576 0,64 3,01556029 0.% 1.72 0,75 2.9841482 0,97 1.96 0,86 2.95338379 0,98 2.44 0,98 2.92324722 0.99 3.8800000000001 1.09 2.89371947 1,02gogogogo90g 0.01 1.53 2.78379903 1.0303157894737 0.05 .54 2.78048954 1,032 0.10 ,56 2.77595182 1,0338823529412 0.15 .58 2.77089774 1.036 0.20 ,60 2.76523386 1.0384 0.25 .63 2.75884272 1.0411428571429 0.30 ,66 2.75157463 1.0443076923077 0.35 .70 2.74323583 1.048 0.40 ,74 2.73357087 1.0523636363636 0.45 .79 2.7222361 1,0576 0.50 2.70875783 1.064 0.55 ,92 2.69246455 5-3 --_ dOWDITI_/_ 1.072 0.60 12.01 2.67237153 1.0822857142857 0.65 12.12 2.64697412 1.096 0.70 12.28 2.61385244 1.1152 0.75 12.49 2,56885068 1.144 0.80 12.81 2,50418031 I. 192 0.85 13,35 2,40334084 1.288 0.90 14.43 2.22420984 1.32 0.91 14,78 2.1702896 1,36 0.92 15.23 2.10645756 1.4114285714286 0.93 15,81 2.02970404 1.48 0.94 16.58 1.9356637 1,576 0.95 17,65 1.81775525 1.72 0.96 19,26 1.66557109 1.96 0,97 21.95 1.46162361 2,44 0,98 27.33 1,1740911 3 8800000000001 O.99 43.46 0.73834595 w = i

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IHC_L"I-/A d.ou_,NG h= 100 ft.

Flight Condition 2 M = 0.85 4.300 rEI/sec Short Period Frequency 0.350 -- Short Period Damping Ratio 2.400 r_/sec Dutch Ro]] Frequency 0.480 -- Dutch Roll Damplng Ratio OmegaD/RollRa_ OmegaD/Omecja$.P. PI (deg/_c) Om_a $. P./Roll Rate 1.169696969697 0.01 24637.1276 0,01 0.05 1,1768421052632 0.03 4927.42552 0,10 1,1866666666667 0.06 2463,71276 1.1976470588235 0,08 1642,47517 0,15 1,21 0,11 1231.85638 0.20 1,224 0,I4 985,485103 0,25 030 1.24 0,17 821.237586 0.35 1.2584615384615 0,20 703.917931 1.28 0,22 615.928189 0.40 1.3054545454545 0,25 547.491724 0,45 1.336 0,28 492.742552 1,3733333333333 0.31 447.947774 0.55 1,42 0,33 410.618793 0.60 1.46 0.36 379.032732 0.65 1,56 0.39 351,958965 1.672 0,42 328.495034 0.75 1,84 0,45 307.964095 0.80 0,85 2.12 0,47 289,84856 0.90 2.68 0,50 273.745862 0.91 2.8666666666667 0,51 270.737666 3.1 0,51 267.794865 0.92 3,4 0.52 264.91535 0.93 0.94 3.8 0.52 262.097102 0.95 4.36 0,53 259.338185 0,96 5,2 0,54 256,636746 0,97 6.6000000000001 0.54 253.991006 0.98 9.4000000000001 0.55 251.399261 17,800000000001 0.55 248.859875 0.99 0,01 0.65 210.628293 1,169696969697 1.1768421052632 0.05 0.66 209.349474 0,10 0.66 207.616244 1,1866666666667 0,15 0.67 205,712755 1.1976470588235 0,20 0.68 203.6i2625 1.21 0,25 0.68 201.283722 1.224 1.24 0.30 0,69 198,6865i3 1.2584615384615 0,35 0.70 195.771796 0,40 0,71 192,477559 i.28 0,45 0.73 188,724515 1.3054545454545 050 0.75 184,409638 i.336 0.55 0.77 179,39656 1.3733333333333

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1.42 0,60 0.79 173.500898 1.48 0.65 0,83 166.467078 1.56 0.70 0.87 157.930305 1.672 O.?S 0,93 147.351241 1.84 0.80 1.03 133.897432 2.12 0.85 1.18 116.212866 2.68 0.90 1,50 91.9295805 2,8666666666667 0.91 1.60 85,9434683 3,1 0.92 1,73 79.4746051 3.4 093 1.90 72.4621399 3.8 0.94 2.12 64.8345463 4.36 0.95 2.43 56.5071733 5.2 0.96 2,90 47,3790915 66000000000001 0.97 3.68 37.3289812 9,4000000000001 0.98 5.25 26,2097102 17,80000000000 1 0.99 9.93 13.8410829

Coup& aNC

H = 0.90 h = 15000 ft, Flight Condition 4 Open L_p O. 140 radlsec Short Per loclFrequency 0.280 -- Short Period Damping Ratio 4.190 r_Isec Dutch RollFrequency 0.370 -- Dutch Roll Damping Ratio PI (deglsec) OmeoaDIRollRate Omega D/OmegeS. P.

Ome_ S. P./Roll Rate 802.139037 1.1046464646465 0.30 0.01 160.427807 0.05 1.1090526315789 1.50 80.2139037 0.10 1.1151111111111 2.99 53,4759358 0,15 1,1218823529412 4,49 40.1069519 0.20 1.1295 5.99 32,0855615 0.25 1,1381333333333 7,48 1,148 898 26,7379679 0,30 r 22,9182582 0.35 1.1593846153846 10.48 20,0534759 1,1726666666667 11,97 04O 17.8253119 0.45 1.1883636363636 13.47 16.0427807 0,50 1.2072 14,96 14,5843461 1.2302222222222 16.46 0.55 13.368984 1.259 17,96 0.60 12.3406006 0,65 1.296 19.45 11.4591291 0.7O 1.3453333333333 20.95 10.6951872 0.75 1.4144 22.45 10.026738 0.80 1.518 23.94 1.6906666666667 25.44 9.43692985 0.85 2.036 26.94 8.91265597 0.90 2.1511111111111 27.24 8,8147147 0.91 8.71890258 2.295 27.53 0.92 8.62515094 0.93 2.48 27,83 8.53339402 0.94 2.7266666666667 28,13 8.44356882 0,95 3,072 28.43 3,59 28.73 8.35561497 0.96 0,97 4,4533333333334 29.03 8.26947461 0,98 6.1800000000001 29.33 8.18509222 8,10241452 0,99 11.36 29,63 7.26150006 1.1046464646465 0,01 33,06 0.05 33,19 7.23265077 1,1090526315789 7,19335526 1,11511II1Iilll 0. I0 33.37 7.14993899 1,1218823529412 0.15 33,58 0.20 33,80 7.10171791 1.1295 0.25 34,06 7.04784768 1.1381333333333 0.30 34,36 6.98727385 1.148 0.35 34.70 6.91866208 1.1593846153846 1.1726666666667 0.40 35,10 6.84029878 1.1883636363636 0,45 35.57 6.749946 6,64462423 1,2072 0,50 36.13 6.52027758 1.2302222222222 0.55 36.82 P_UL- EORC_a4 E-R!

1.259 0.60 37,68 6,37123938 1.296 0.65 38.79 6.18934442 1.3453333333333 0.70 40.26 5.96238135 1.4144 0.75 42.33 5.67123188 1,518 0.80 45.43 5,28418338 1.6906666666667 0.85 50.60 4.74451323 2.036 0.90 60.93 3,93977916 2.1511111111111 0.91 64.38 3.72895214 2.295 0.92 68.69 3.4951592 2.48 0.93 74.22 3.2344316 2.7266666666667 0.94 81.61 2.94183021 3.072 0,95 91.94 2.61112968 3.59 0,96 107.44 2.23437058 4.4533333333334 0.97 133.28 1.80121041 6.1800000000001 0.98 184.96 1.29795961 11.36 o.g9 33999 0.70610831 i

P 4 uu Bo

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h --15000 ft.

Flight Condition 4 M = 0.90 3.800 rMlsec Short Period Frequency 0.350 -- Short PeriodDamping Ratio 4.180 red/_c DutchRollFrequencu 0.380 -- Dutch Roll Damping Ratio Pl(_glsec) Omega 5. P./RollRate Omega D/RollRate Omega D/Omega S. P.

21772.3453 0.01 1.1343434343434 0.01 1.14 0.06 4554.46906 0.05 1.1477777777778 0.11 2177.23453 0.10 0.15 I.I 564705882353 0,I7 1451.48969 0.20 1.16625 0.22 1088.61727 870.893812 0.25 1.1773333333333 0.28 725.744843 0.30 1.19 0.33 622.067009 0.35 1.2046153846154 0.39 544.308633 O40 1.2216666666667 0.44 1.2418181818182 0.50 483.829896 0.45 435.446906 0.50 1.266 0.55 O rr 1 2955555555556 0.6 i 395.860824 1.3325 0.66 362.872422 0.60 O.b.., 1.38 0.72 334.959158 0,70 1.4433333333333 0.77 311.033504 L 075 1.532 0.83 290.297937 0.80 1.665 0.88 272.154316 w 256.145239 0.85 1.8866666666667 0.94 241.914948 0.90 2.33 0.99 2.4777777777778 1.00 239.256542 0.91 2.6625 1.01 236.655927 0.92 2.9 1.02 234.11124 0.93 0.94 3.2166666666667 1.03 231.620695 229.182582 0.95 3.66 1.05 4.325 1.06 226.795264 0.96 5.4333333333334 1.07 224.457168 0.97 I 0.98 7.6500000000001 108 222.166789 0.99 14.3 1.09 219.92268 1.1343434343434 0.01 1.25 191.937862 0.05 125 190.985485 1.14 1.1477777777778 O.10 1.26 189.691295 0.15 1.27 1.1564705882353 188.265448 0.20 1.28 186.686776 1.16625 0.25 1.30 184.92932 1.1773333333333 1.19 0.30 1.31 182.960885 0.35 1.33 180.741053 1.2046153846154 1.2216666666667 0.40 1.34 178.218379 1.2418181818182 0.45 1.37 175.326353 1.266 0.50 1.39 171.977451 12955555555556 0.55 1.43 168.054123 w

A-E & 747.-

.5"--- _'.'5 -- .9o

Ft.- I _ ft'l" CO/,/b I"1' I b#J L)- I _usP /P , _-J._/_s_ 1.3325 _0.6_00' 1.47 163.394711 1.38 0.65 1.52 157.770618 1.4433333333333 0.70 1.59 150.847658 1.532 0.75 1.69 142.117136 1.665 0.80 1.83 130.764837 1.8866666666667 0.85 2.08 115.401124 2.33 0.90 2.56 93.4435421 2.4?777??777??8 0.91 2.73 87.8704519 2.6625 0.92 2.93 81.7740669 2.9 0.93 3.19 75.0770528 3.216666666666? 0.94 3.54 67.6860476 3.66 0.95 4.03 59.4872822 4.325 0.96 4.76 50.3406828 5.4333333333334 0.97 5.98 40.0718012 7.6500000000001 0.98 8.42 28.4605821 14.3 0.99 15.73 15.2254163 w = --= M = 1.60 h = 30000 ft.

Flight Condition 7 Open Loop 2.160 radlsec Short Pemod Frequency' 0.345 -- ,Short PeriodDamping Ratlo 3,150 radlsec DutchRollFrequency 0.120 -- Dutch RollDamping Ratio

pl (oeg/sec)

OmegaD/RollRate OmegaD/OmegaS. P.

Om_a $. P./Roll Rate 12375.8594 1.0418181818182 0,01 0.01 2475.17189 0.05 1.0435789473684 0.07 1237.58594 0.I0 1.046 0.15 825.057296 0,15 1.0487058823529 0.22 618.792972 0,20 1.05175 0.29 495.034377 0.25 1.0552 0.36 412.528648 1.0591428571429 0.44 1.0636923076923 0.51 353,595984 0,35 309,396486 1.069 0,58 0.40 275,0}9099 1,0752727272727 0.66 0.45 247.517189 1,0828 0,73 0.50 225,01 _ 56c6 1.092 0.80 0.55 206.264324 1,1035 0,88 0.60 190.397837 0.65 1.1182857142857 0.95 1,138 1,02 I76.797992 0.70 165,011459 1,1656 1,09 0.75 154,698243 1,207 1.17 0.80 145.598346 1,276 1.24 0.85 137.509549 1.414 1.31 135.998455 1,46 1.33 0,91 134.520211 1,5175 1.34 0.92 133,073757 0.93 1,5914285714286 1.36 1.69 1,37 131,658079 0.94 1.828 1.39 130.272205 0.95 2,035 1.40 128.915202 0.96 2.38 1.41 127,58618 0.97 3,07 1.43 126.28428 0.98 0.99 5.1400000000001 1.44 125.008681 118.790972 1.0418181818182 0.01 1.52 0.05 1.52 118.590543 1.0435789473684 118.316056 1.046 0.I0 1.53 118.010775 1,0487058823529 0.15 1.53 0.2O 1.53 117.669213 1,05175

0,25 1.54 117,28449

1.0552 116,847877 1.0591428571429 0,30 1.54 0,35 1.55 116.348114 1,0636923076923 i 115.770434 0,40 1.56 1,069 115,095074 1,0752727272727 0.45 1,57 114.294971 1.0828 0.50 1,58 0,55 1,59 113,332046 1,092 B oR c b_-_ C O N O I "TI ON rT_ OP£t# zoo PF I. 1035 0.60 1.61 112.150969 I. 1182857142857 0.65 1.63 !10.668135 1.138 0.70 1.66 108.750962 I, 1656 0.75 1.70 106.17587 1.207 0.80 1,76 102.534047 1.276 0.85 1.86 96.989494 1.414 0.90 2,06 87.5237584 1.46 0.91 2.13 84.7661605 1.5175 0.92 2,21 81.5542632 1.5914285714286 0.93 2.32 77.7657236 1.69 0.94 2.46 73.2299375 1.828 0.95 2.67 67.701638 2,035 0.96 2.97 60.8150341 2.38 0.97 3.47 51.9994094 3.07 0.98 4.48 40.3122457 5 1400000000001 0.99 7.50 240775475 Co U @1..._ tv_ c h_r-.oU_-"rl oU' S

., FC 7

M = 1.60 h = 30000 ft.

Flight Condition ?

1.850 radlsec Short Period Frequency 0.300 -- Short Period Damping Ratio 1.650 r_l/sec Dutch Roll Frequency 0.100 -- Dutch Roll Damping Ratio Pl(_glsec) OmegaS. P./Roll Rate Omega D/RollRate Omega D/Omega 5. P.

10599.6944 0.01 1.030303030303 0.01 2119.93888 0.05 1.0315789473684 0.04 1059.96944 0.10 1 0333333333333 0.09 1.0352941176471 0.13 706.646295 0.15 0.20 1.0375 O. 18 529.984721 0.25 1.04 0.22 423.987777 353.323147 030 1.0428571428571 0.27 302.848412 0.35 1.0461538461538 0.31 1.05 0.36 264.992361 0.40 1.0545454545455 0.40 235.548765 0.45 1.06 0.45 211.993888 &- 0.50 1.0666666666667 0.49 192.721717 0.60 1.075 0.54 176.661574 163.072222 0.65 1.08571 428571 43 0.58 151.424206 0.70 1.1 0.62 14t.329259 0.75 1.12 0.67 1.15 0.71 152.49618 0.80 1.2 0.76 124.702287 0.85 117.774382 0.90 1.3 0.80 1.3333333333333 0.81 116.480158 0.91 1375 0.82 115.21407 0.92 0.93 1.4285714285714 0.83 113.975209 I12.762707 0.94 1.5 0.84 111.575731 0.95 1.6 0.85 1.75 0.86 110.413464 0.96 2 0.87 109.2752 0.97 0.98 2.5 0.87 108.160147 107.06762 0.99 4.0000000000001 0.88 O.01 0.92 102.879387 1.030303030303 0.05 0.92 102.75214 1.0315789473684 1.0333333333333 O.10 0.92 102.577688 1.0352941176471 O.15 0.92 102.383412 0.20 0.93 102.165729 1.0375 0.25 0.93 101.920139 1.04 1.0428571428571 0.30 0.93 101.640905 1.0461538461538 0.35 0.93 101.320608 i 100.949471 1.05 0.40 0.94 100.514344 1.0545454545455 0.45 0.94 1.06 0.50 0.95 99.9971172 1.0666666666667 0.55 0.95 99.3721352

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The purpose of this appendix is to show the calculations for the landing gem" sizing and the horizontal stabilator actuator sizing as discussed in Chapter 11.

Appendix 4: Table of Contents page w w w L._._o_'_ C,_A_. Cj, I,_N _ 2"3 MaC, _o V I_ F'r C c_ _ HOYEr4 Fs 44S ?LoNF, q _oo_.

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APPENDIX 5 The purpose of this appendix is to present calculations to determine the es_atcd increase in drag due to the open internal weapons bay and to estimate the maximum yawing moment that is induced by the firing of the Vulcan cannon.

Ap_ndix 5: Table of Contents w page w I T- F(_I G_'T C ¢> _,,I ;_ i't-, o N 1_ --- i._,

H : _0.ooo _¢

va- _94.-_ r_A

C = o.ooo_m _lug/_ _

llm <4'¢

/_ = _.,oe • ,o "_ ib.s,c/_, z

WW_ ZZZ ooc sq_ L = 1_6 in W: q;n I.

(,3.1o_. ,o "_ )

'7

R_L = S. 9 2 4-, to

w

%- o.4ss (lo:_ _. yz.s_

e_= o.,%S(io_=_._z4x,C_ -z's'

w

(_ o "= O.OOZZ93

_ rl C_= O. o09z T J Ac._ 0 /qO_aLl'_E T¢_ R_Ee_NCqr t_Rw_ f- i : I m

_',:, = (.0.oo,_tX_,_.,/_,rT. _')

d O-- O. Oo05Z -- -,',oo C%= O. O4O_B w

o._,oosz (._oo,_

0.040_, =-- w w

f

I ¢¢-- [ ( i RE_ 12. t r_ N i

I

J2..EI:'. 1"2.. 7..

"70 _.Ps ..,.- h w w

I