Hypersonic Vehicle Systems Integration Vehicle Propulsion Integration and Force Accounting

1
Hypersonic Vehicle Systems IntegrationVehicle
Propulsion Integration and Force Accounting

• 12 September, 2007
• Dr. Kevin G. Bowcutt
• Senior Technical Fellow
• Chief Scientist of Hypersonics
• Boeing Phantom Works

2
Propulsion System Design Integration
Challenging Due to Aero-Propulsion Requirements
and Complex Flow Physics

• Inlet Isolator
• Prevents combustion-induced flow separation from
  migrating forward to inlet
• Flow processed similar to a normal shock

Angled exhaust flow generates lift and pitching
moment

• Inlet
• Increases air pressure and reduces Mach number
  for efficient combustion
• Goal is to minimize overall entropy rise of
  propulsion process
• Shocks must be weak enough to avoid massive
  boundary layer separation
• Boundary layer transition critical

Spilled air generates lift pitching moment

• Nozzle
• Expands engine flow to extract work
• Frozen chemical reactions rob thrust from engine

• Combustor
• Adds thermal energy to flow
• Pressure rise, hence thrust, decreases with Mach
  number
• At low Mach, pressure rise can separate inlet
  flow (causing inlet unstart)
• At low Mach, heat addition can choke flow
  (requires downstream injection)
• Complete fuel mixing and combustion challenging

Aerodynamic testing very challenging for highly
integrated engine-airframe configurations -
Propulsion interaction effects force accounting
issues
3
Hypersonic Vehicle Propulsion Requirements and
Considerations
Combustor
Forebody and Inlet
Nozzle and Aftbody

• Forebody
• Compression ratio
• Inlet
• Stability
• Spillage
• Forebody-Inlet Interaction
• Flow separation
• Flow distortion
• Boundary layer transition
• Lateral flow spillage

Combustor-inlet interaction Flow distortion Skin
friction Heat transfer and cooling Fuel-air mixing
Combustor-nozzle interaction Free steam
interactions Skin friction and heating Lift and
trim effects

• General

• Inlet
• Pressure recovery

• Subsonic

Thermal choking
Ground effects

• Inlet
• Start
• Pressure recovery
• Shock interactions
• Combustor interaction and isolation

Ignition delay Flame holding Thermal choking Mode
transition

• Over expansion and flow
• separation
• Drag
• Lift and trim effects
• Thermal choking

• Supersonic

• Forebody
• Mass capture
• Contraction ratio
• Compression efficiency
• Inlet
• Shock on cowl lip Mach
• Shock interactions
• Heat transfer and cooling

Turbulence modeling Chemical Kinetics Shock
losses / injector drag Mixing efficiency

• Balancing losses
• Divergence
• Chemical kinetics
• Under expansion
• Friction
• Viscous interactions
• Lateral flow effects
• Heat transfer and cooling

• Hypersonic

4
Approach to Aero-Propulsion Force Accounting
Critical for Performance Analysis Ease, Accuracy
and Consistency

• Force Accounting Options
• Cowl-To-Tail
• Often referred to as cycle performance
• Hard to avoid force discontinuities due to use of
  different tools for aero and propulsion analysis,
  with different modeling approaches, assumptions
  and fidelity for each
• Ramp-To-Tail
• Include inlet ramps in propulsion forces and
  moments
• Works well for low-speed performance analysis
  where its difficult to include propulsion
  flowpath surfaces in aero analysis
• Can also lead to force discontinuities
• Nose-To-Tail
• A consistent approach, since the forebody
  contributes significantly to propulsion
  performance, particularly at hypersonic speeds
  (avoids force discontinuities)
• Difficult to implement at low speeds because the
  forebody generally cannot be removed from aero
  analysis force and moment integrations

5
Momentum Theory Extremely Valuable for Propulsion
Analysis and Accurate Force Accounting

• Newtons 2nd Law of Motion applied to a continuum
  fluid
• Draw a control volume, S, around a body, or
  through an engine flowpath from free stream to
  engine exit

6
First-Order Vehicle Propulsion Requirements

• Distinct propulsion systems are likely required
  for feasible and efficient and operation in
  different flight regimes (but may require close
  integration, i.e., combined cycles, to achieve
  needed synergy between propulsion systems)
• Subsonic through supersonic
• Hypersonic
• Exo-atmospheric
• Engine specific impulse or TSFC (i.e., engine
  efficiency) and thrust-to-weight ratio are
  primary propulsion parameters that drive design
• Fuel selection is a key driver of vehicle
  performance, size and cost
• Engine sizing is critical for achieving
  sufficient vehicle performance
• T/Wvehicle 0.33 for takeoff
• T/Wvehicle 0.25 for transonic and any other
  thrust pinch-points
• T/D 3 across entire hypersonic flight regime

7
Hypersonic Air-breathing Propulsion Requires High
Dynamic Pressure Flight to Generate Adequate
Thrust
350
Legend
300
SSTO Flight Corridor Air-breathing Flight Corridor
250
SSTO Descent
Dynamic Pressure (psf)
200
Altitude (x1000 FT)
1000
150
2000
3000
100
SSTO Ascent
50
0
15
10
5
0
30
25
20
Mach Number
8
Engine Specific Impulse and Thrust-to-Drag Ratio
Determine Propellant Consumption

• From Newtons Second Law F (mV), an
  equation can be derived for propellant fraction
  required between any two trajectory points
• This equation can be used todetermine thrust /
  drag (T / D)requirements for reasonablepropella
  nt mass fractions
• For realistic ?p (0.3 - 0.4),(T/D) 3-4 required

? Wfuel Winitial
V2 2
V2 2
- ? ( ) g?h gV Isp 1 - (D/T)
1 - exp ,
where ? g?h ?E
- ?E Qf ?p 1 - (D/T)
1 - exp ,
since ?p gVIsp / Qf
Governing Parameters For Single Stage-To-Orbit Vo
26,000 ft/sec, ho 500,000 ft Q 51,600
BTU/lb, q? 1,000 lb/ft2
1.0
Requirement
.9
?
.8
.2
.7
MF/MT
.6
.3
.5
.4
.4
.6
.3
1.0
.2
.1
0
1.5
2
3
4
5
6
7
8
9
10
20
30
100
50
T/D
From Jones and Donaldson 3
9
Numerous Propulsion Technology Options Exist for
Hypersonic Vehicles

• Turboramjets, ramjets/scramjets, rockets,
  combined-cycle engines

10
Typical Propulsion System Speed Limits
From Raymer, Daniel P., Aircraft Design A
Conceptual Approach, Fourth Edition, AIAA
Education Series, 2006
11
Low Speed Engine Choice Driven by Isp,
Thrust-to-Weight System Volume Tradeoffs

• Several low speed engine concepts are viable
  candidates
• Turbojet
• Turbo-ramjet
• Rocket
• Air augmented rocket (e.g., RBCC)
• Liquid air cycle engine
• Low speed engine generally sized for transonic
  pinch-point
• Large scramjet inlet and nozzle create high
  transonic drag

Payload Fraction
10,000
0.6
8,000
Vehicle Thrust Loading
0.7
Specific Impulse (Sec)
6,000
0.8
0.9
4,000
1.0
Payload Fraction
2,000
0.20 0.10 0.0
0
0
10
20
30
40
50
60
70
80
90
100
Engine Thrust / Weight Ratio
Specific Impulse and Thrust-to-Weight Can be
Traded
From Curran, et al. 10
12
Low-Speed Engine Example SR-71 J-58 Engine
Operation Versus Subsonic Mach Number
13
Low-Speed Engine Example SR-71 J-58 Engine
Operation Versus Supersonic Mach Number
14
Turbine-Based Combined-Cycle (TBCC) Engine Mode
Transition Challenge

• TBCC engines comprised of turbines mounted over
  ramjets/scramjets are promising propulsion
  systems for hypersonic cruise aircraft and
  reusable launch vehicles
• Feasibility of mode transition between engines
  has not yet been fully established
• Aerodynamic, mechanical and thermal interactions
  must be understood and managed
• Thrust margin must remain adequate with little or
  no operability risk
• Airframe-engine system must be able to tolerate
  and control events that could cause mission
  failure or vehicle loss (e.g., inlet unstart,
  engine flameout, thermal transients, etc.)

15
Other TBCC Engine Challenges

• Thermal issues of Mach 3-4 turbine engines
• Lightweight, durable, high-temperature materials
• High-temperature bearings, bushings, seals and
  wear surfaces
• Thermal management during operation and after
  shutdown
• Transonic and high Mach thrust performance
• Thrust pinch points/sufficient thrust over entire
  speed range
• Transonic nozzle drag download of large
  scramjet nozzle
• Ground testing verification and validation
• Facility size and Mach number/enthalpy
  constraints to accommodate both flowpaths of a
  TBCC engine
• Integrating turbine inlets and nozzles in complex
  dual-mode scramjet flowpath shapes
• Strong aerodynamic interactions between highly
  integrated dual inlets and dual nozzles

16
Subsonic and Supersonic Turbine Engine
Performance Can Be Estimated via Simulation Tools

• NASA Glenns EngineSim 1.7a can be used to
  estimate turbofan, turbojet and ramjet engine
  weight, thrust and efficiency, as a function of
  speed and altitude (http//www.grc.nasa.gov/WWW/K-
  12/airplane/ngnsim.html)
• Engine size, and component design, material and
  performance parameters can be varied to match
  existing engines

17
Supersonic Combustion Ramjets (scramjets) Provide
for Efficient Hypersonic Propulsion

• Shock waves compress, heat slow flow
• Flow remains supersonic throughout engine
• Fuel cooling capacity limits endothermic
  hydrocarbon use to Mach 8 cryo CH4 has higher
  limit cryo H2 can be used to highest Mach
  desired, though will eventually exceed
  stoichiometric fuel-air ratio for cooling

Performance levels verified in flight by X-43A
18
Vehicle Forebody Considerations Driven by
Requirements of Different Speed Regimes

• High Speed
• Inlet mass capture
• Compression efficiency
• Optimum inlet contraction ratio
• Boundary layer transition
• Lateral flow spillage
• Inlet flow distortion
• Flow separation
• Corner flow effects
• Aerodynamic heating insulation/cooling
  requirements
• Shock interactions with leading edges
• Optimum inlet axial location
• Longitudinal vehicle stability
• Pitch / yaw sensitivity

• Low Speed
• Flow spillage drag
• Inlet starting
• Flow separation
• Inlet stability (unstart)
• Compression efficiency
• Combustor isolation

19
Forebody Requirements and Design Options

• Requirements
• High compression efficiency
• Low drag and high mass capture (high wair / D)
• Flow uniformity to inlet
• Minimal crossflow
• Uniform boundary layer transition
• Good volumetric efficiency
• Maintain aerodynamic center aft for stability
• Design Options
• Cones
• Wedges
• Ogives
• Spatular bodies
• Bielliptic cross-sections (lenticular)
• Waveriders
• Inward turning / Busemann inlets

Spatular Bodies
0.08
Opt. Spatular Body
0.06
CD
0.04
2b
0.02
2a
0
0
0.2
0.4
0.6
0.8
1.0
b a
Drag of conical and optimum nose shapes with
elliptical bases of constant area
Minimum drag shapes
From Townend 11
Stream surfaces carved from compression flow
fields
20
Hypersonic Engine Thrust-to-Weight Can Be a
Function of Vehicle Forebody Type

• Assumptions
• Thrust capture area
• Engine length equal for all cases
• Engine weight engine internal wetted area
  inlet perimeter
• Engine T/W capture area / inlet perimeter

Axisymmetric
2-D Planar
Inward-Turning
?
R
h
?
R
W
H
H
R
w
W

• Capture area (Ac)
• Inlet area (Ai)
• Contraction ratio (CR)
• Ac/Ai
• Inlet height (H)
• For equal CR
• Inlet width (W)
• For equal Ac
• Inlet perimeter (p)

R2 ? ? RH R / (2H) h/2 2w 2H 2W h
4w
?/2
?/2
R2 H2 (R / H)2 Rh h CR 2w h/R
2w / CR 2H W 2h CR 2w / CR
Rw hw R / h h w 2h 2w
?/2
CR - 1 CR - CR
P gt P2D for h lt 2w ? T/W lt (T/W) 2D
P lt P2D for h lt w ? T/W gt (T/W) 2D
21
Inlet Air Capture Requirement VariesDramatically
With Vehicle Mach Number

• Shock-on-lip at design (maximum) Mach number
• 14 cone external compression (drawn to scale)

Thrust-to-Drag Ratio Requirements Drive Need for
Increasing Air Capture With Mach Number To
Achieve Sufficient Integrated Vehicle Performance
Courtesy of Scott Halloran, Pratt Whitney
Rocketdyne
22
Capture Area Drives Thrust-to-Drag RatioCapture
Area Therefore a First-Order Configuration Driver

• The rocket equation be used to estimate vehicle
  capture area requirement, hence relative engine
  size
• Solve for A? / Axs from equation for T/D

T wair fs ? Isp? q? A? (2g fs ? Isp?) / (M?
?RT?)
Axs
where fs 0.02916 for H2 fuel and Isp?
nose-to-tail engine specific impulse
D q? Cp (Axs - A?) q? Cf (Awet - Awetfp)
where (Awet - Awetfp) (Axs - A?) 11 / tan2
?c for a cone
A?
Axs total vehicle frontal area and A?
flowpath capture area
2g fs ? Isp? M? ?RT? (T/D)
Axs / A? / Cp
Cf 1 1 / tan2 ?c 1
Capture Area (A? / Axs)

• Assumptions
• Cp for 9 cone angle
• Drag-due-to-lift and trim drag ignored
• q? 2,000 psf trajectory
• Shock-on-cowl-lip for all cases
• But approximately valid for undersped inlet
  conditions

A? / Axs ()
M?
Cp
Cf
Isp?
2
3
4
T / D
2 5 10
0.088 0.062 0.055
0.025 0.015 0.001
3800 3500 1800
6 10 26
8 14 35
11 17 42
23
Inlet Design Tradeoffs Driven By Inlet
Performance, Drag, Weight and Complexity

• Inlet design options include external, internal
  and mixed compression concepts
• Each has different performance, drag, weight and
  complexity characteristics

External Compression Inlets
Internal Mixed Compression Inlets
From Raymer, Daniel P., Aircraft Design A
Conceptual Approach, Fourth Edition, AIAA
Education Series, 2006
24
Inlet Type and Shock Number Trends With Maximum
Mach Number
From Raymer, Daniel P., Aircraft Design A
Conceptual Approach, Fourth Edition, AIAA
Education Series, 2006
25
Supersonic (Turbine) Inlet Design Layout Approach

• Select capture area size
• Select desired number of ramps and ramp angles
• At design Mach number, position ramps so that all
  ramp shocks focus on the cowl leading edge
• Another option is not focusing the shocks at the
  design condition, but then there will always
  exist supersonic spillage drag for Mach numbers
  at least as high as the design Mach number
• Note For fully variable ramps and cowl,
  spillage drag can be made zero for all Mach
  numbers, but will likely result in increased cowl
  drag at supersonic speeds

26
Turbine Inlet Pressure Recovery Estimation

• Turbine inlet pressure recovery can be estimated
  from MIL standard curve (MIL-E-5008B), or trends
  from existing analytical predictions or actual
  airplane inlet data

From Raymer, Daniel P., Aircraft Design A
Conceptual Approach, Fourth Edition, AIAA
Education Series, 2006
27
Inlet Performance and Operational Issues /
Requirements at Hypersonic Speeds

• High compression efficiency
• Inviscid (turn angles and shock strengths)
• Friction and heat loss (pressure distribution,
  boundary layer transition and wetted area)
• Viscous phenomena may affect performance and
  operability
• Entropy swallowing effects on boundary layer
  transition
• Shock / boundary layer interactions (potential
  flow separation)
• Corner flows
• Note High air total temperature prevents use of
  boundary layer bleed to
• compensate for viscous effects.
• Actively cooled surfaces
• Required when convective heating rates exceed
  threshold
• Leading edge heating and drag
• Edges must be blunt to reduce heating and
  accommodate active cooling
• Shock / leading edge interactions increase
  heating dramatically
• Optimum axial location
• Influences inlet capture and / or performance,
  vehicle trim and c.g.

28
Hypersonic/Scramjet Inlet Design Layout Approach
Fuselage Reference Line (FRL)
hc, Ac

• Select a cowl lip axial station 0.45 0.55
  times body length is reasonable
• Select desired number of inlet ramps, and whether
  cowl and/or ramps are variable
• At design Mach number, position cowl vertically
  so that forebody shock is on the cowl leading
  edge, then position ramps so that all ramp shocks
  focus on the cowl lip
• Another option is to not focus the shocks at the
  design condition, but then there will always
  exist supersonic spillage drag for Mach numbers
  at least as high as the design Mach number

29
Total Flow Turning in Optimum 3-Shock, 4-Shock,
and Isentropic Inlets
p4 1/2 atm for q8 2000 psf flight
Van Wie, David M., Purdue University Short
Course, NASA Marshall Space Flight Center, Feb.
7-11, 2000
30
Accounting for All Sources of Inlet Drag Critical
to Accurate Force Accounting
Throat Bleed Drag
Secondary Airflow Drag
Additive Drag (pressure integral on captured
streamtube)

• Inlet Spillage Drag Additive Incremental Cowl
  Drag (or Thrust) due to spilling flow
• Bypass Drag (due to flow momentum loss)
• Bleed Drag (due to flow momentum loss)
• Secondary Airflow Drag (due to flow momentum
  loss)

From Raymer, Daniel P., Aircraft Design A
Conceptual Approach, Fourth Edition, AIAA
Education Series, 2006
31
Effect of Forebody Boundary Layer Transition on
Inlet Performance
1.0
0.98
All Laminar Flow For M? ? 20
Kinetic Energy Efficiency ?KE
0.96
Legend
Laminar Transition to Turbulent at Re? / M 290
0.94
0.92
14
12
10
8
6
20
18
16
22
Mach Number
From Harsha 9
32
High-Speed Propulsion Requirements Impact
Transonic Aero Performance

• Inlet drag
• Body ramps required for high speed inlet
  efficiency, but . . .
• Ramps produce high spillage drag at transonic and
  low supersonic speeds
• Nozzle drag
• Large base area required for high speed thrust,
  but . . .
• Low nozzle pressure ratios result in aftbody flow
  separation, drag and large pitching moments (trim
  drag)

33
Transonic Inlet Drag

• The forebody ramp and inlet system of
  air-breathing hypersonic vehicles typically
  generate large drag at transonic conditions
• Large ramps required for adequate and efficient
  high-speed compression
• Inlet throat area must be small at high speeds to
  meet contraction ratio requirements for good
  engine performance
• Impact Maximum throat size at low speeds is
  limited by mechanical (variable geometry)
  constraints and fuel penetration across flowpath
  duct

Typical Transonic Inlet Spillage Phenomenon
Subsonic Spillage Zone
Pressure
x
34
Nozzle Considerations vs. Speed Regime

• Hypersonic
• Optimum balance of losses
• Divergence
• Friction drag
• Underexpansion
• Finite-rate chemistry (kinetics)
• Freestream interactions (vertical and lateral)
• Heat transfer and cooling
• Combustor flow profile effects
• Viscous interactions
• Relaminarization
• Contribution to vehicle lift and trim

• Subsonic/Transonic
• Overexpansion and drag (internal and external)
• Drag reduction
• Freestream interactions
• Combustor interactions(e.g., thermal choke
  location and orientation)
• Flow separation
• Ground effects

35
Nozzle Expansion Area Requirement Varies
Dramatically With Mach Number
Internal Thrust Control Volume
Vehicle
A9
p4 p0
Design
Ideal Nozzle Expansion Area to Base Area Ratio
vs. Mach
Cowl Hinge
Flap
Slip Line
Vo
V10
4
9
10
6
Design Point Operation
5
4
Ideal Expansion Area / Vehicle Base Area
p4 p0
A9
3
lt Design
2
1
Overexpanded Operation
0
0
2
4
6
8
10
12
14
16
18
20
Freestream Mach
A9
p4 p0
gt Design
General trend shown actual values are vehicle
design specific
Underexpanded Operation
From Heiser Pratt 15
Optimum Nozzle Size is a Tradeoff Between Vehicle
Drag, Nozzle Overexpansion and Underexpansion
Losses All a Function of Body Cross-Sectional
Area
36
Transonic Nozzle Drag Can Be Quite Large

• Nozzle region of space plane designs typically
  generate large drag at transonic conditions
• Engines produce relatively low nozzle pressure
  ratios
• Large areas required for high Mach scramjet
  efficiency
• Confirmed by experiment

Sym
NPR
Sym
NPR
0.00 2.02 2.53 3.06 3.57 4.06
0.00 1.31 2.02 2.53 3.08 3.61 4.00
-3.0
-3.0
0.0
0.0
Cp
Cp
Re 7 X 106
Re 7 X 106
3.0
3.0
Centerline Static Pressure Distribution, Mach
0.60, 20-Degree Afterbody
Centerline Static Pressure Distribution, Mach
0.60, 10-Degree Afterbody
From Novak and Cornelius 12
37
Over-expanded Nozzle Base Drag First-Order
Analysis Approach

• First determine fraction of nozzle expansion area
  filled by engine exhaust flow (i.e., area after
  flow expands to ambient pressure)
• Apply base pressure to remaining nozzle area

38
External Burning An Approach To Transonic
Nozzle Drag Reduction

• Injecting and burning fuel external to the engine
  cowl can fill nozzle base region with hot, low
  density gases at ambient pressure to reduce or
  eliminate nozzle drag

Constant Pressure Combustion Control Volume
Constant Pressure Mixing Control Volume
p ? p?
Normal, Sonic H2 Injection
Engine Exhaust Stream
Nozzle Expansion Surface
From Trefny 13
39
Scramjet Nozzle Efficiency and Loss Mechanisms

• Optimum nozzle performance requires a careful
  balance between loss mechanisms across entire
  Mach range
• Low divergence losses (e.g., perfect nozzle)
  result in excessive viscous losses due to long
  lengths
• Gradual expansion to reduce kinetic losses
  results in large viscous and divergence losses
• Typical values of losses at hypersonic speeds

Underexpansion Losses
Viscous Losses (friction drag and heat transfer)
Pe gt Po
Divergence Losses (Non-parallel flow, 2-D and 3-D)
?v
H OH M ? H2O M H H M ? H2 M
V
Chemical Kinetic Losses (Dissociated H2O and H2)
-? CFG 1/2 - 1 0 - 1/2 1 - 2 1/2 - 1
Loss Mechanism Divergence Underexpansion Friction
Kinetics
40
Nozzle Design Variables Can Be Used to Tailor
Nozzle Thrust, Lift and Pitching Moment
Lateral Expansion
Trailing Edge Angle
Longitudinal Curvature
Body Flap Angle
Initial Expansion Angle
Nozzle Height
Chordal (Avg.) Angle
Lateral Curvature
Side-Wall Fences
Cowl Angle (Fixed or Variable)
Cowl Length
Variables Can Also Be Used To Minimize Thrust
Losses and / or Maximize Vehicle Performance
Through Multidisciplinary Design Optimization
(MDO)