Hypersonic Flight Vehicles

J. Gordon Leishman

53 Hypersonic Flight Vehicles

Introduction

Hypersonic flight is generally defined as speeds at or beyond Mach 5. However, unlike what happens as the aircraft approaches Mach 1, there are no distinctive changes in the flow characteristics as an aircraft reaches hypersonic Mach numbers. Such high Mach numbers have only been achieved by rockets, re-entry spacecraft, and special high-altitude research aircraft. Indeed, every spacecraft that re-enters the Earth’s atmosphere will travel at hypersonic speeds. Theodore von Kármán was one of the first to analyze the problems of hypersonic flight, stating: “At such speeds, even in rarified air, the surface will be heated to the temperatures none of the known materials can withstand. The problem of the thermal barrier is much more complicated than the problem of the sound barrier.”

By necessity, hypersonic flight vehicles look quite different from supersonic flight vehicles. The reason is that when flight vehicles begin to fly at hypersonic Mach numbers, many additional aerodynamic and other technical challenges occur, a summary of which is shown in the figure below. The main issues include the proximity of the shock waves to the vehicle, kinetic heating, airframe/engine flow interactions, viscous interactions, chemical reactions, and surface erosion. The shapes suitable for hypersonic flight then become highly specialized, including a tightly integrated propulsion system.

Very complex aerodynamics and other issues occur in hypersonic flight.

In more detail, these effects at hypersonic flight Mach numbers include, but are not limited to, the following:

1. The Mach angle is small, so the shock waves bend back to steep angles and can begin to interact with the airframe, disrupting the flow in the boundary layer. Usually, the boundary layers are relatively thin, even laminar. Still, the resulting flow field becomes very complex at hypersonic speeds as the shock waves touch the boundary layer, and flow separation is likely.

2. The shock waves are very strong at hypersonic speeds, and the air undergoes significant temperature changes that affect its chemistry. The need to include chemistry in any aerodynamic analysis of hypersonic flows significantly raises the complexity of the needed predictive methods.

3. The proximity of shock waves to the body and the effects of skin friction cause a large amount of aerodynamic heating on vehicles flying at hypersonic speeds. Even high-temperature aluminum and titanium, commonly used for supersonic airplane construction, will melt.

4. The flow over the vehicle can strongly interact with the propulsion system. In this regard, the propulsion system design must be tightly integrated into the entire vehicle design. Usually, a hypersonic vehicle must be designed so that its shape pre-compresses the air before it enters the engine.

5. The sonic booms from aircraft flying at Mach > 3 and hypersonic speeds will be loud enough to likely damage structures on the ground and permanently damage human ears. It is unlikely that a hypersonic flight vehicle of any significant size could be allowed to fly over populated areas. Nevertheless, this technical challenge will remain in the future.

Hypersonic technologies may also be poised to revolutionize military systems by enabling vehicles that travel above Mach 5 with high maneuverability, making them nearly impossible to intercept. Platforms like hypersonic glide vehicles and cruise missiles offer rapid global strike capability but present significant thermal protection, guidance, and propulsion challenges. Several countries are investing heavily in these systems, prompting parallel efforts in counter-hypersonic defense using advanced sensors and interceptors. On the civil side, hypersonic flight remains a long-term ambition, with research targeting Mach 5 passenger aircraft that could drastically reduce intercontinental travel times. However, significant hurdles remain in developing propulsion systems, heat management, and regulatory approval. Sonic booms from hypersonic flight vehicles will be extreme, and a need for boom mitigation introduces further constraints in their design.

Learning Objectives

Appreciate the unique aerodynamic and thermal challenges faced by aircraft operating at hypersonic speeds, including strong shock interactions, high aerodynamic heating, and real gas effects.
Know how to apply analytical techniques such as Newtonian and shock-expansion theories to determine aerodynamic coefficients on bodies in hypersonic flow conditions.
Develop an understanding of the role of high-temperature gas dynamics, boundary layer behavior, thermal protection strategies, and air-breathing engine concepts, such as scramjets, in the design and analysis of hypersonic vehicles.

History

The challenges associated with hypersonic flight first became evident during the early Cold War period, when post-WWII ballistic missile development pushed vehicles into Mach 5 and higher regimes. To achieve intercontinental range, these missiles had to exit the atmosphere and re-enter it near the target location. Initial tests, based on highly pointed German V-2 designs, revealed catastrophic thermal failure during re-entry. The slender nose shapes experienced intense aerodynamic heating and shock wave compression that led to complete structural failure, often melting entirely during descent. These events underscored the extreme aerodynamic and thermal loads in hypersonic flight.

The answer to this problem, which was research led by H. Julian Allen & Alfred Eggers at the NACA, was opposite to that expected based on conventional aerodynamic wisdom, which was to make the nose of the missile much more rounded, as shown in the schlieren visualization images below. This “blunt body” design created a shock wave that stood off the nose and caused a boundary layer to flow over the nose cone’s surface, preventing the excessive heat of re-entry from melting the structure. In contrast, pointed nose shapes allowed the shock wave to touch the surface with a boundary layer that was too thin to offer much thermal protection. Tests with the modified missiles confirmed the validity of this method, although the nose still needed to be made of high-temperature resistant materials.

NACA research showed that at hypersonic Mach numbers, a rounded nose created a shock wave that stood off the surface, reducing the surface temperatures and preventing the body from melting.

The information gained about high-speed flight and hypersonics in the wind tunnel and through flight tests contributed significantly to the development of the Mercury, Gemini, and Apollo space flight programs. Flight test work on hypersonics was first conducted by the U.S. Air Force and NASA during the 1950s. The X-15 research airplane flew for nearly ten years from 1959 and set the world’s unofficial speed and altitude records of 4,520 mph (Mach 6.7) and 354,200 feet during a program to investigate all aspects of piloted hypersonic flight.

The X-17 was a missile program that used a three-stage rocket to get the nose section to reach conditions between Mach 10 and 20 in the lower atmosphere. The nose cones survived briefly before vaporizing, but long enough for valuable measurements to be made. Ultimately, the solution to minimizing kinetic heating effects on re-entry vehicles was also to use ablative materials to help shed the heat buildup.

During the 1980s, NASA began considering a single-stage hypersonic flight vehicle to replace the Space Shuttle. The National Aerospace Plane (NASP) was intended to take off from a standard runway used by airliners. Once the aircraft had reached sufficient airspeed, scramjet engines would power the aircraft into hypersonic flight. Finally, a rocket system would take the NASP into orbit. The NASP eventually became the Rockwell X-30 research vehicle. While the project did not reach the flight stage, a one-third-scale demonstrator, the X-43, was built. Three were built, the first being destroyed after an in-flight malfunction, but the other two flew successfully with the scramjet operating for approximately 10 seconds.

Since the X-43A’s flights in the early 2000s, hypersonic development has accelerated worldwide. The U.S. X-51A Waverider achieved sustained scramjet-powered flight at Mach 5.1 for over 200 seconds, while Stratolaunch’s reusable Talon-A vehicle reached hypersonic speeds and completed recovery missions by 2025. Venus Aerospace demonstrated a rotating detonation rocket engine for its Mach 9-capable Stargazer platform, and Australia’s Hypersonix developed the hydrogen-fueled SPARTAN scramjet for the Delta-Velos vehicle, designed for speeds up to Mach 12. Meanwhile, Switzerland’s Destinus and China’s hypersonic aircraft efforts aim toward commercial applications, with India also entering the arena through successful hypersonic missile tests. Together, these advances signal growing global investment in reusable, high-speed flight technologies across defense, research, and future commercial transport.

Hypersonic Flow Modeling

Hypersonic aerodynamics introduces physical effects beyond those encountered in subsonic or supersonic regimes. As flight speeds exceed Mach 5, the flows are increasingly dominated by strong shock waves, extreme post-shock temperatures, thick boundary layers, and significant thermal loads. The physical complexity of hypersonic flow makes detailed analysis challenging, especially in early-stage design or conceptual studies. Strong shocks, entropy^[1] gradients, viscous interactions, and chemistry effects contribute to this difficulty. To this end, engineers often rely on simplified models that approximate the surface pressures and aerodynamic forces without solving the complete set of governing flow equations, which, for the most part, are impractical.

Shocks at High Mach Numbers

In supersonic flow, a wedge-shaped body generates an oblique shock wave at its leading edge, as shown in the figure below. Unlike a normal shock, which stands perpendicular to the flow, an oblique shock forms at an angle relative to the surface, allowing the flow to remain supersonic downstream. As the freestream Mach number increases, the shock angle decreases, bringing the shock closer to the surface.

Shock waves and streamlines over a 20° half-angle wedge at Mach 2 and 20. (Adapted from Anderson, 2000.)

At hypersonic speeds, this shock standoff distance becomes very small, and the region between the shock and the body surface is often referred to as a shock layer. In hypersonic flow over slender bodies, this shock layer can be extremely thin, particularly at high altitudes where the Reynolds number is low. Under such conditions, the shock layer may interact with or merge into the boundary layer, forming a fully viscous shock layer. This complex regime significantly influences heat transfer, surface pressures, and overall aerodynamic characteristics.

Slender Bodies

When a supersonic flow encounters a sharp compression surface, such as a wedge or ramp, the flow is deflected through an angle $\theta$ relative to the incoming stream, as shown in the figure below. In response, an oblique shock forms at an angle ${\beta}$ with respect to the upstream flow direction. The relationship between the deflection angle $\theta$ , the shock angle ${\beta}$ , and the upstream Mach number $M_1 = M_\infty$ is described by the $\theta$ – ${\beta}$ – $M$ relation, a fundamental expression in compressible flow theory. As discussed in the previous chapter, this relationship is given by

(1) $\begin{equation*} \tan\theta = 2 \cot\beta \frac{M_1^2 \sin^2\beta - 1}{M_1^2(\gamma + \cos 2\beta) + 2} \end{equation*}$

where $\gamma$ is the ratio of specific heats, which is 1.4 for air.

For a given Mach number and deflection angle, this relation admits two possible solutions: a weak shock and a strong shock. The weak solution is typically observed in external aerodynamic flows, as it produces lower pressure and temperature increases and allows the flow to remain supersonic downstream of the shock. Although mathematically valid, the strong solution is usually associated with internal or confined flows and is rarely seen in freestream conditions.

At high Mach numbers and small deflection angles, the shock angle ${\beta}$ is only slightly greater than the deflection angle $\theta$ , resulting in a thin shock layer that hugs the body. However, as the deflection angle increases toward a critical value, the oblique shock becomes increasingly intense and eventually can no longer remain attached to the surface. Beyond this point, the shock detaches, forming a curved bow shock ahead of the body. This shock detachment phenomenon is significant in hypersonic flow, where small changes in body geometry can dramatically alter surface pressure and thermal loads. Understanding the onset of shock detachment is essential for designing vehicles operating in these regimes.

Blunt Bodies

In contrast to slender bodies that produce oblique shocks, blunt bodies generate normal shock waves when the incoming supersonic flow is brought to rest abruptly, typically at a stagnation point. This occurs, for instance, at the nose of a reentry vehicle, such as shown in the figure below, where the flow cannot turn smoothly around the body as it does over a wedge. Instead, the air must decelerate to near-zero velocity over a very short distance. In supersonic or hypersonic regimes, the only physically admissible way for this to occur is through the formation of a detached bow shock standing off from the surface. This strong normal shock induces a sudden and severe rise in pressure and temperature, initiating the thin but highly energetic shock layer that envelops the front of the vehicle.

A shock wave forms around the nose of a blunt reentry vehicle, which at the centerline forms a normal shock.

Notice that the shock is normal only on the centerline, where the freestream stagnates directly. It is oblique elsewhere, wrapping around the body and becoming more inclined with increasing distance from the centerline. Across an oblique shock, only the velocity component normal to the shock is compressed, while the tangential component remains unchanged. The governing equations for pressure, temperature, and density changes are still those of a normal shock, but they apply to the normal component of the upstream Mach number, i.e.,

(2) $\begin{equation*} M_{n,1} = M_1 \sin \beta \end{equation*}$

where $M_1$ is the freestream Mach number and ${\beta}$ is the local shock angle between the flow and the shock front. The downstream Mach number is computed from the turned flow, and both components are accounted for. The flow behind the shock remains supersonic if the shock is weak (i.e., associated with a small flow deflection), or subsonic if the shock is strong (large deflection).

A normal shock wave abruptly increases pressure, temperature, and density, accompanied by a sharp reduction in Mach number. As the upstream Mach number $M_1$ increases, these thermodynamic changes become more pronounced. In the strong-shock limit ( $M_1 \gg 1$ ), the post-shock pressure and temperature scale approximately with ${M_1^2}$ . For a calorically perfect gas, the leading-order approximations for the temperature and pressure ratios across a normal shock are

(3) $\begin{equation*} \frac{T_2}{T_1} \approx \frac{2(\gamma - 1)}{(\gamma + 1)^2} M_1^2 \quad \text{and} \quad \frac{p_2}{p_1} \approx \frac{2\gamma}{\gamma + 1} M_1^2 \end{equation*}$

where $\gamma$ is the ratio of specific heats, which is 1.4 for air. These relations are valid as long as the post-shock temperature remains within the range where the gas can be considered calorically perfect, generally below 1,000–1,500 K for air.

These expressions highlight the sensitivity of post-shock conditions to the freestream Mach number. Even moderate increases in $M_1$ can significantly increase downstream pressure and temperature. In hypersonic flow, where Mach numbers are very high, normal shocks generate steep thermal gradients that strongly influence boundary layer behavior and surface heating. A key consequence of this heating is the dramatic rise in convective heat flux experienced by the surface of a hypersonic vehicle. Empirically, the convective heat flux $q$ scales approximately as the cube of the freestream velocity, i.e.,

(4) $\begin{equation*} q \ \propto \ \varrho_{\infty} \, V_{\infty}^3 \end{equation*}$

Given that the velocity is related to the Mach number by $V_{\infty} = M_{\infty} \, a_{\infty}$ , and the speed of sound is $a_{\infty} = \sqrt{\gamma \, R \, T_{\infty}}$ , assuming constant ambient density and temperature leads to the simplified scaling

(5) $\begin{equation*} q \ \propto \ M_{\infty}^3 \end{equation*}$

This result implies that doubling the Mach number results in an approximate eightfold increase in convective heating. Such sensitivity underscores the importance of accurate thermal protection system design in hypersonic flight regimes.

Several important limiting cases illustrate this behavior. At the stagnation streamline, ${\beta} = 90^\circ$ , so $M_{n,1} = M_1$ , and the result is a full normal shock. As the shock angle decreases with increasing distance from the centerline, $M_{n,1}$ becomes smaller and the shock weakens. In the limit as ${\beta} \to 0^\circ$ , the shock becomes a Mach wave. Consequently, the strength of the shock, and therefore the heating and pressure rise, varies continuously over the body’s surface. The highest heating and pressure occur near the nose, where the shock is nearly normal and the freestream kinetic energy is most effectively converted into internal energy. This has direct consequences for aerodynamic loading and thermal protection system design.

Shock Layers and Boundary Layer Behavior

As previously described, at hypersonic flow speeds, the bow shock formed ahead of a blunt body detaches and stands off from the surface, enclosing a narrow region of compressed gas known as the shock layer, as shown in the figure below. For hypersonic speeds, this shock layer can be quite thin. Although thin relative to the body’s overall size, this shock layer contains steep pressure, temperature, density, and velocity gradients. As the flow passes through the bow shock, it decelerates, compresses, and heats rapidly, especially along streamlines near the stagnation region. The balance between shock curvature and surface shape governs the geometry of the shock layer. Near the nose, where the shock is nearly normal, the post-shock flow slows significantly, and the layers are relatively thick. Away from the stagnation region, the shock becomes increasingly oblique, the post-shock velocity remains higher, and the shock layer narrows.

Shock layer over a blunt body with a normal shock at the nose and an oblique shock off-axis. The boundary and entropy layers form at the stagnation point.

When the surface deflection angles are small and the freestream Mach number is high, the local flow can often be modeled using the thin shock-layer approximation. This approach assumes that variations in flow properties occur primarily in the direction normal to the surface, rather than along it, i.e.,

(6) $\begin{equation*} \left| \frac{\partial p}{\partial n} \right| \gg \left| \frac{\partial p}{\partial s} \right| \end{equation*}$

where $n$ is the surface-normal direction, and $s$ is tangential. This simplification reduces the governing equations, making estimating surface pressure and thermal loads easier.

Near the stagnation point, a useful scaling for the shock layer thickness is

(7) $\begin{equation*} \delta_s \ \propto \ \frac{R_n}{M_\infty^2} \end{equation*}$

where $R_n$ is the nose radius of curvature and $M_\infty$ is the freestream Mach number. As the Mach number increases, the shock layer becomes thinner and more compressed, a defining feature of hypersonic flow.

In this regime, the viscous boundary layer also behaves differently than in subsonic or low-supersonic flows. It grows more rapidly, driven by strong shock compression, elevated surface temperatures, and increased gas viscosity. As a result, the boundary layer becomes dynamically significant and may interact with shock waves, affecting the overall flow structure.

Unlike lower-speed flows, where boundary layers tend to transition quickly to turbulence, hypersonic boundary layers often remain laminar for longer distances. An approximate scaling for the laminar boundary layer thickness $\delta$ over a flat plate at hypersonic speeds is

(8) $\begin{equation*} \frac{\delta}{L} \ \propto \ \frac{1}{\sqrt{Re_x}} \left(1 + \frac{\gamma - 1}{2} M_\infty^2 \right)^{1/2} \end{equation*}$

where $L$ is a characteristic length, $Re_x$ is the Reynolds number based on distance from the leading edge, $\gamma$ is the specific heat ratio, and $M_\infty$ is the freestream Mach number. In the high-Mach limit, this simplifies to

(9) $\begin{equation*} \delta \ \propto \ \frac{M_\infty}{\sqrt{Re_x}} \quad \text{or} \quad \ \delta \ \propto \frac{M_\infty^2}{\sqrt{Re_x}} \end{equation*}$

depending on wall temperature and thermal boundary conditions. These scalings highlight how boundary layer thickness increases more rapidly with Mach number than in incompressible flows, making it a dominant factor in aerodynamic performance and heating.

The rise in surface pressure in high-speed flow is commonly expressed using the pressure coefficient, i.e.,

(10) $\begin{equation*} C_p = \frac{p_w - p_\infty}{\tfrac{1}{2} \, \varrho_\infty \, V_\infty^2} \end{equation*}$

where $p_w$ is the wall (surface) pressure and $q_\infty = \tfrac{1}{2} \varrho_\infty V_\infty^2$ is the freestream dynamic pressure. At the stagnation point, where the flow comes to rest and the shock is strongest, the pressure coefficient reaches its maximum value, denoted $C_{p,0}$ . This value can be estimated using normal shock relations from the post-shock pressure $p_2$ .

In regions with steep surface deflection, such as the nose of a blunt body, the incoming flow behaves like a direct impact. In this limit, the pressure coefficient approaches the Newtonian approximation

(11) $\begin{equation*} C_p(\theta) = C_{p,0} \cos^2 \theta \end{equation*}$

where $\theta$ is the angle between the freestream direction and the surface normal. This model, while approximate, provides a simple and effective way to estimate surface pressure distributions in hypersonic flow over blunt geometries.

Force on a Reentry Cone

Consider a slender reentry vehicle modeled as a cone with half-angle $\theta$ , moving at hypersonic speed through the atmosphere. Using the impact theory, the pressure coefficient on the cone surface is approximated as

(12) $\begin{equation*} C_p(\theta) = 2 \sin^2 \theta \end{equation*}$

The surface pressure becomes

(13) $\begin{equation*} p(\theta) = C_p(\theta) \, q_\infty = 2 q_\infty \sin^2 \theta \end{equation*}$

A differential ring element at axial position $x$ has radius $r = x \tan \theta$ and surface area

(14) $\begin{equation*} dA = 2 \pi x \tan \theta \, \frac{dx}{\cos \theta} \end{equation*}$

The axial component of the pressure force is

(15) $\begin{equation*} dF_x = p(\theta) \cos \theta \, dA = 2 q_\infty \sin^2 \theta \cos \theta \, 2 \pi x \tan \theta \, \frac{dx}{\cos \theta} = 4 \pi q_\infty x \tan \theta \sin^2 \theta \, dx \end{equation*}$

Integrating over the cone from $x = 0$ to $x = L$ , the total axial force is

(16) $\begin{equation*} F_x = \int_0^L 4 \pi q_\infty x \tan \theta \sin^2 \theta \, dx = 2 \pi q_\infty \tan \theta \sin^2 \theta \, L^2 \end{equation*}$

Using the impact theory, this result gives the total axial force on a slender cone, expressed in terms of dynamic pressure $q_\infty$ , cone length $L$ , and cone half-angle $\theta$ . It provides a useful estimate of reentry drag based on an idealized surface pressure distribution.

Check Your Understanding #1 – Analysis of the nose of a reentry vehicle

Consider a reentry vehicle traveling at a freestream Mach number of $M_\infty$ = 10 through the upper atmosphere at an altitude of approximately 40 km. The vehicle has a hemispherical nose with a radius of curvature $R_n$ = 0.5= m. At this altitude, the freestream conditions are: $\varrho_\infty$ = 0.003996 kg/m $^3$ , $T_\infty$ = 250 K, and $p_\infty$ = 2,871 Pa. Assume the gas behaves as a perfect gas with a specific heat ratio of $\gamma$ = 1.4 and gas constant $R$ = 287 J/(kg K).

Show solution/hide solution.

This example focuses on the leading nose geometry of the vehicle, where hypersonic compression effects and shock-layer behavior are most intense. The surface pressure and pressure coefficients can be predicted at key locations along the nose using the thin shock-layer approximation and Newtonian theory. The isentropic relation gives the freestream velocity for a perfect gas, i.e.,

$V_\infty = M_\infty \sqrt{\gamma R T_\infty}$

Substituting the values gives

$V_\infty = 10 \sqrt{1.4\, 287\, 250} = 3,170 \text{ m/s}$

The dynamic pressure is calculated using

$q_\infty = \frac{1}{2} \rho_\infty V_\infty^2 = \frac{1}{2} \, 0.003996 \, (3170)^2 \approx 2.01 \times 10^4 \text{ Pa}$

To estimate the pressure at the stagnation point, the normal shock approximation for high Mach numbers can be used, i.e.,

$\frac{p_2}{p_\infty} = \frac{2\gamma}{\gamma + 1} M_\infty^2 - \frac{\gamma - 1}{\gamma + 1} = \frac{2.8}{2.4} \, 100 - 0.167 = 116.5$

so that

$p_w = 116.5 \, p_\infty \approx 3.34 \times 10^5 \text{ Pa}$

The stagnation-point pressure coefficient is then

$C_{p,0} = \frac{p_w - p_\infty}{q_\infty} = \frac{3.34 \times 10^5 - 2871}{2.01 \times 10^4} = 15.5$

To estimate the pressure at an off-stagnation location, say $\theta = 45^\circ$ , the Newtonian pressure approximation can be used, i.e.,

$C_p(\theta) = C_{p,0} \cos^2 \theta = 15.5 \, \cos^2(45^\circ) = 15.5 \, 0.5 = 7.75$

The local surface pressure is

$p_w(\theta) = p_\infty + C_p(\theta) \, q_\infty$

so that

$p_w(45^\circ) = 2871 + 7.75 \, 2.01 \times 10^4 \approx 1.58 \times 10^5 \text{ Pa}$

Finally, the stagnation-point shock-layer thickness can be estimated using

$\delta_s \ \propto \ \frac{R_n}{M_\infty^2} = \frac{0.5}{100} = 0.005 \text{ m} = 5 \text{ mm}$

These results demonstrate key features of hypersonic flow over blunt reentry geometries, including strong compression and thin shock layers.

Entropy Gradients and Flow Interactions

An increase in entropy typically means energy is lost to irreversible processes such as shock compression and viscous dissipation in aerodynamic flows. Across a shock wave, entropy always increases because of the irreversible nature of the compression process. Across a curved bow shock, entropy increases significantly, especially near the stagnation streamline. This behavior creates a high-entropy layer near the body centerline, displacing lower-entropy gas outward. The result is a distortion in the outer flow and the surface boundary layer, which are relatively thick in hypersonic flows.

A viscous interaction parameter is used to characterize the influence of viscous effects in hypersonic flow, particularly in the coupling between the boundary layer and the outer inviscid field. The thick boundary layer in hypersonic flow can exert a major displacement effect on the inviscid flow outside the boundary layer, causing a given body shape to appear much thicker than it is. Two related but distinct forms are commonly used, namely $\chi$ and $\bar{\chi}$ . The parameter $\chi$ is defined using surface heat transfer as

(17) $\begin{equation*} \chi = \frac{q_w}{\varrho_\infty \, V_\infty^3} \end{equation*}$

where ${q_{w}}$ is the wall heat flux, $\varrho_\infty$ is the freestream density, and $V_\infty$ is the freestream velocity. This dimensional form quantifies the relative magnitude of energy extracted from the flow by conduction at the wall. A larger value of $\chi$ indicates stronger viscous interaction and greater boundary-layer displacement effects.

In contrast, $\bar{\chi}$ is a nondimensional scaling parameter defined in terms of freestream properties and Reynolds number as

(18) $\begin{equation*} \bar{\chi} = \frac{M_\infty^3}{\sqrt{Re_L}} \end{equation*}$

where $M_\infty$ is the freestream Mach number and $Re_L = \varrho_\infty V_\infty L / \mu_\infty$ is the Reynolds number based on a characteristic length $L$ and freestream viscosity $\mu_\infty$ . This form of the interaction parameter is often used to classify flow regimes by the strength of viscous interaction. Values of $\bar{\chi}$ less than one indicate weak interaction, values near unity indicate moderate interaction, and larger values of $\bar{\chi}$ greater than 3 correspond to strong interaction with significant effects on shock shape, boundary layer growth, pressure gradients, as well as lift and drag.

Boundary Layer Scaling and Surface Heating

By far the most dominant aspect of high temperatures in hypersonic flow about a body is the resultant high heat-transfer rates to the surface. The kinetic energy of a high-speed flow is dissipated by the influence of friction within a boundary layer. The extreme viscous dissipation that occurs within hypersonic boundary layers can create very high temperatures. The boundary layer in hypersonic flow is generally thicker than in subsonic or transonic conditions because of compressibility and elevated viscosity. For a laminar boundary layer over a flat plate, the thickness scales as

(19) $\begin{equation*} \delta \ \propto \ \frac{x}{\sqrt{Re_x}} = \frac{x}{\sqrt{\dfrac{\varrho_\infty \, V_\infty \, x}{\mu}}} \end{equation*}$

where $x$ is the distance from the leading edge and $\mu$ is the dynamic viscosity. In hypersonic conditions, temperature-dependent viscosity leads to additional thickening, particularly under cold-wall assumptions.
Thermal loads are a primary design constraint in hypersonic vehicles. Near the stagnation point, the kinetic energy of the freestream is converted into internal energy, resulting in intense surface heating. A representative scaling law for stagnation-point heat flux is

(20) $\begin{equation*} q_{\text{stag}} \ \propto \ \varrho_\infty^{1/2} \, V_\infty^3 \end{equation*}$

highlighting the strong cubic dependence on velocity. Even modest speed increases can cause dramatic rises in heat flux, necessitating robust thermal protection. For adiabatic^[2] surfaces, the recovery temperature indicates the effective surface temperature from frictional heating. For a laminar boundary layer, the recovery temperature is approximately

(21) $\begin{equation*} T_r = T_\infty \left( 1 + \frac{\gamma - 1}{2} M_\infty^2 \right) \end{equation*}$

assuming a recovery factor near unity. Heat is transferred into the surface if the wall temperature is below this value.

High-Temperature Gas Behavior

At post-shock temperatures above 2000 K, air exhibits “real gas” or high-temperature behavior. Initially, vibrational modes of diatomic molecules such as $\text{O}_2$ and $\text{N}_2$ become energetically active. As the temperature increases further, these molecules dissociate into atoms. At even higher temperatures, ionization occurs, producing a partially ionized gas.

These processes increase the internal energy storage capacity of the gas and alter its thermodynamic behavior. The specific heats become functions of temperature, and the perfect gas assumption no longer applies. In some situations, the gas can be treated as being in chemical equilibrium; in others, finite-rate chemical reactions lead to nonequilibrium effects, especially within the boundary and shock layers. These high-temperature phenomena influence heat transfer rates, boundary layer structure, and surface pressure distributions. They become especially important in high-Mach cruise, atmospheric reentry, or any flight regime where post-shock temperatures exceed several thousand K.

Hypersonic “Newtonian” Theory

Like all flight vehicle design aspects, the ability to predict quantitatively aerodynamic forces for the right (physical) reason is fundamental. To this end, a relatively parsimonious but validated method must often be used, providing the necessary focus to establish initial designs from which more detailed analysis can follow. A widely used approximation in hypersonic aerodynamics is Newtonian impact theory. Although it does not account for boundary layer effects, viscous heating, or “real gas” (high-temperature) behavior, it effectively estimates surface pressures based on local flow deflection. The core idea is geometric and kinematic: the freestream flow is treated as a stream of particles that transfer momentum normal to the surface upon impact. This leads to a pressure distribution that depends only on the local surface inclination angle relative to the freestream.

Isaac Newton did not know about shock waves or supersonic or hypersonic flows. Still, in his Principia of 1687, he hypothesized some simple hydrodynamic principles, which have seen application to hypersonic flows in recent decades. Newton proposed that fluid flow forces exerted on a surface consisting of a uniform stream of “particles” in which the time rate of change of the normal component of the momentum of each particle would be transferred to the plate, whereas the tangential component would be preserved. He assumed that the upper surface always remains in the “shadow” of the flow and is hence uninfluenced by the freestream flow.

The figure below shows a hypersonic flow impinging on a thin body, resulting in a shock wave close to the surface and a thin shock layer. In a supersonic flow, disturbances cannot be propagated upstream, so there is no streamline curvature as the flow approaches the wing. Indeed, the resulting flow field looks remarkably like that proposed by Newton more than three centuries ago, despite the physical disparity in detail. Indeed, although the so-called Newtonian aerodynamics theory is flawed for several reasons, it is a reasonably good model to describe the lift and drag of bodies in hypersonic flows.

The hypersonic flow about a body looks remarkably similar to the flow model assumed by Isaac Newton over three centuries ago.

In the context of hypersonic lift and drag estimation, Newtonian theory can be applied to slender flat plates or symmetric wedges at an angle of attack. The central assumption is that only the windward surface contributes significantly to pressure forces. At the same time, the leeward side is effectively pressure-less, which is a valid approximation in high Mach number regimes. The surface pressure coefficient is given by

(22) $\begin{equation*} C_p = 2 \sin^2 \theta \end{equation*}$

where $\theta$ is the angle between the local surface normal and the freestream direction.

Consider a symmetric wedge of half-angle $\varepsilon$ flying at an angle of attack $\alpha$ . Newtonian theory can provide an approximate method for estimating the aerodynamic pressure distribution in hypersonic flow. The inclination angles of the upper and lower surfaces relative to the freestream are $\theta_u = \varepsilon - \alpha$ and $\theta_\ell = \varepsilon + \alpha$ , respectively. According to Newtonian theory, the local pressure coefficients on each surface are

(23) $\begin{equation*} C_{p,u} = 2 \sin^2(\varepsilon - \alpha) \quad \text{and} \quad C_{p,\ell} = 2 \sin^2(\varepsilon + \alpha) \end{equation*}$

These pressure forces act normal to the wedge surfaces and are decomposed into lift and drag components by projection onto the freestream direction. Assuming unit span and chord length ${c}$ , the total lift and drag coefficients are given by

(24) $\begin{equation*} C_L = \frac{1}{2} (C_{p,\ell} - C_{p,u}) \cos \alpha. \quad \text{and} \quad C_D = \frac{1}{2} (C_{p,\ell} + C_{p,u}) \sin \alpha \end{equation*}$

These expressions highlight the nonlinear dependence of aerodynamic forces on both the angle of attack and the wedge geometry. The theory is most accurate for slender wedges at high Mach numbers, where pressure forces dominate over viscous effects and shock-layer interactions. Despite its simplicity, Newtonian theory remains useful for preliminary hypersonic design estimates.

Consider also a flat plate at an angle of attack $\alpha$ , where in this case the surface is uniformly inclined at $\theta = \alpha$ . Applying Newtonian theory and resolving the normal pressure force into freestream-aligned components, the resulting aerodynamic coefficients are

(25) $\begin{equation*} C_L = 2 \sin^2\!\alpha \cos\alpha \quad \text{and} \quad C_D = 2 \sin^3\!\alpha \end{equation*}$

The corresponding lift-to-drag ratio is

(26) $\begin{equation*} \frac{C_L}{C_D} = \frac{2 \sin^2\!\alpha \cos\alpha}{2 \sin^3\!\alpha} = \frac{\cos\alpha}{\sin\alpha} = \cot \alpha \end{equation*}$

While the Newtonian “hypersonic” theory is elementary, at least in terms of predicting the lift and drag coefficients, the question is how well it might work in practice, i.e., can it be used to predict anything useful? To this end, the results shown in the figures below compare the Newtonian theory with NASA wind tunnel measurements on a hypersonic body at Mach numbers of 6 and 10; measurements are available for angles of attack of -5.5^o to 35^o at Mach numbers from 1.5 to 10.

Predictions using “Newtonian” hypersonic theory against wind tunnel measurements.

Despite its simplicity, the Newtonian theory is in remarkable agreement with the measurements. The drag is underpredicted because the Newtonian theory does not account for wave drag and viscous effects. However, the drag equation can be modified empirically to

(27) $\begin{equation*} C_D = 2 \, \sin^3\!\alpha + C_{D_{w}} + C_{D_{0}} \end{equation*}$

where ${C_{D_{w}}}$ is the wave drag and ${C_{D_{0}}}$ is the viscous shear stress drag, both of which will depend on the specific body shape. One approximation used for the wave drag of a body in hypersonic flow is

(28) $\begin{equation*} { C_{D_{w}} = \dfrac{2}{\gamma \, M_{\infty}^2} } \end{equation*}$

which is attributed to Nonweiler. Notice that for $M_{\infty}$ = 10 then $C_{D_{w}}$ = 0.02. Furthermore, it is often assumed that high-speed bodies have a viscous friction drag proportional to their wetted area times a known constant, i.e., the skin friction coefficient, $C_f$ , which comes from boundary layer theory or the von Driest equation, i.e.,

(29) $\begin{equation*} C_f \approx \frac{0.074}{Re^{~0.2}} \left( \frac{1}{1 + 0.15 \, M_{\infty}^2} \right)^{0.58} \end{equation*}$

Therefore, approximately, for the upper and lower surfaces, then

(30) $\begin{equation*} C_{D_{0}} = 2 C_f \end{equation*}$

Naturally, there are flow interference effects to consider, too. In this case, a value of $C_{D_{w}} + C_{D_{0}} = 0.09$ aligns well with the wind tunnel measurements. The corresponding lift-to-drag ratio now becomes

(31) $\begin{equation*} \frac{C_L}{C_D} = \frac{2 \, \sin^2\!\alpha \, \cos\alpha}{2 \, \sin^3\!\alpha + C_{D_{w}} + C_{D_{0}}} \end{equation*}$

which again is seen to be in good agreement with the wind tunnel measurements. Notice that in the absence of viscous effects, the Newtonian theory becomes singular at exactly zero angle of attack, but this is of no practical significance.

It can be concluded that, despite its limitations, Newtonian “hypersonic” aerodynamics can still be used in some instances to provide reasonable estimates and initial approximations for aerodynamic lift and drag. Therefore, it can be helpful in conceptual design studies or in situations where a rapid assessment of the aerodynamic characteristics of a hypersonic vehicle is required.

Check Your Understanding #2 – Hypersonic flow over a slender cone

A sharp, slender cone with a half-angle of $\theta$ = 10 $^\circ$ is flying at Mach $M_\infty$ = 10 through the upper atmosphere. The freestream conditions are: air density $\varrho_\infty$ = 0.0185 kg/m $^3$ , static temperature $T_\infty$ = 270 K, ratio of specific heats $\gamma$ = 1.4, and gas constant $R$ = 287 J/kg/K. The cone has a base diameter of $D$ = 0.5 m. Assume the surface is smooth and inviscid, and that Newtonian impact theory applies uniformly across the surface. Using Newtonian theory, derive an expression for the surface pressure distribution as a function of the cone angle. Integrate this distribution to find the total axial force (drag) in terms of $q_\infty$ , $L$ , and $\theta$ . Then derive the corresponding drag coefficient $C_D(\theta)$ using the base area as the reference. Numerically evaluate the freestream velocity, dynamic pressure, cone length, total axial force, and drag coefficient for the given conditions.

Show solution/hide solution.

The pressure coefficient on the surface of a slender cone using Newtonian theory is

$C_p(\theta) = 2 \, \sin^2 \theta$

The surface pressure becomes

$p(\theta) = C_p(\theta) \, q_\infty = 2 \, q_\infty \, \sin^2 \theta$

A differential ring element at axial position $x$ has radius $r = x \, \tan \theta$ and surface area

$dA = 2 \, \pi \, x \, \tan \theta \, \frac{dx}{\cos \theta}$

The axial component of the pressure force is

$dF_x = p(\theta) \, \cos \theta \, dA = 2 \, q_\infty \, \sin^2 \theta \, \cos \theta \, 2 \, \pi \, x \, \tan \theta \, \frac{dx}{\cos \theta} = 4 \, \pi \, q_\infty \, x \, \tan \theta \, \sin^2 \theta \, dx$

The total axial force is obtained by integration over the cone length, i.e.,

$F_x = \int_0^L dF_x = 4 \, \pi \, q_\infty \, \tan \theta \, \sin^2 \theta \int_0^L x \, dx = 2 \, \pi \, q_\infty \, L^2 \, \tan \theta \, \sin^2 \theta$

The reference area is the base area of the cone, i.e.,

$A = \frac{\pi \, D^2}{4} = \pi \, L^2 \, \tan^2 \theta$

The drag coefficient is, therefore,

$C_D = \frac{F_x}{q_\infty \, A} = \frac{2 \, \pi \, q_\infty \, L^2 \, \tan \theta \, \sin^2 \theta}{q_\infty \, \pi \, L^2 \, \tan^2 \theta} = \frac{2 \, \sin^2 \theta}{\tan \theta}$

Substituting $\theta = 10^\circ$ gives $\sin(10^\circ) = 0.1736$ and $\tan(10^\circ) = 0.1763$ , so that

$C_D = \frac{2 \, (0.1736)^2}{0.1763} \approx \frac{0.0603}{0.1763} = 0.342$

The freestream speed of sound is

$a_\infty = \sqrt{\gamma \, R \, T_\infty} = \sqrt{1.4 \, \times \, 287 \, \times \, 270} = 328 \, \text{m/s}$

and so the freestream velocity is

$V_\infty = M_\infty \, a_\infty = 10 \, \times \, 328 = 3280 \, \text{m/s}$

The dynamic pressure is

$q_\infty = \frac{1}{2} \, \varrho_\infty \, V_\infty^2 = \frac{1}{2} \, \times \, 0.0185 \, \times \, (3280)^2 \approx 99{,}450 \, \text{Pa}$

The cone length is

$L = \frac{R}{\tan \theta} = \frac{0.25}{\tan(10^\circ)} \approx 1.418 \, \text{m}$

Finally, the total drag force is

$F_x = 2 \, \pi \, q_\infty \, L^2 \, \tan \theta \, \sin^2 \theta = 6.67 \, \text{kN}$

Other Methods

To obtain more accurate predictions for hypersonic flow about various bodies, researchers and engineers rely on more advanced approaches such as shock-expansion theory, computational fluid dynamics (CFD), and wind tunnel testing. CFD methods consider the complexities and non-idealities of hypersonic flows, including shock wave interactions, boundary layer effects, chemical reactions, and high-temperature gas dynamics. If properly conducted, wind tunnel testing represents the datum or benchmark against which CFD and other methods can be tested and validated.

Shock Expansion Theory

Shock-expansion theory is another widely used method for modeling hypersonic flow. It is based on the observation that the pressure distribution on a slender body in high-speed flow can often be approximated by treating the surface as a series of compression and expansion segments. When the flow turns into the body, a shock wave forms, producing a sudden increase in pressure, temperature, and density. When the flow turns away from the body, an expansion fan forms, causing a smooth and continuous decrease in these properties.

In hypersonic flows, the analysis proceeds by treating each surface segment individually. Compression regions are analyzed using the oblique shock relations, and expansion regions are treated using the Prandtl–Meyer function. The local pressure on the surface can be computed by applying these relations at each surface segment. Once the pressure distribution over the body is known, the aerodynamic forces such as lift and drag can be obtained by integrating the pressure components along the body surface.

Shock-expansion theory is especially practical for slender bodies at high Mach numbers, where the flow remains attached and the interaction between shock and expansion waves is weak. It is commonly applied in analyzing sharp-nosed wedges, cones, and other faceted geometries where curvature effects are minor and the flow can be considered piecewise planar. However, it does not account for viscous effects and viscous-inviscid interactions that characterize many hypersonic flows.

Unsteady Aerodynamics

Piston theory is widely used to analyze unsteady aerodynamic loading in high-speed flows, especially where detailed solutions of the compressible flow equations are impractical. Applications include predicting panel flutter on aircraft skins and missile bodies, aeroelastic modeling of slender wings and control surfaces at high Mach numbers, unsteady pressure estimation for dynamic stability analysis, and preliminary design analysis in hypersonic vehicle concepts.

In supersonic and hypersonic flows, the piston theory simplifies the problem by treating each element of the surface as a one-dimensional piston pushing into the flow, generating pressure waves that propagate downstream. The approximation is valid in regimes where the Mach number is high, the flow remains attached, and the disturbances are of small amplitude.

Consider a thin surface immersed in a uniform supersonic flow of freestream Mach number $M_\infty$ , pressure $p_\infty$ , density $\varrho_\infty$ , and sound speed $a_\infty$ . Let $w(x,t)$ denote the normal velocity of the surface at location $x$ and time $t$ . The first-order, or linear, piston theory approximation gives the surface pressure as

(32) $\begin{equation*} \frac{p - p_\infty}{\varrho_\infty \, a_\infty} = \frac{2}{M_\infty} \frac{\partial w}{\partial t} \end{equation*}$

This expression shows that the unsteady pressure perturbation is proportional to the time rate of change of the surface normal velocity. The simplification assumes small disturbance theory, which is valid when the normal velocity is much smaller than the freestream speed and the flow direction is nearly aligned with the surface.

To account for weak nonlinear effects, a second-order piston theory expression can be used, i.e.,

(33) $\begin{equation*} \frac{p - p_\infty}{\varrho_\infty \, a_\infty} = \frac{2}{M_\infty} \frac{\partial w}{\partial t} + \left( \frac{\gamma + 1}{4} \right) \left( \frac{2}{M_\infty} \frac{\partial w}{\partial t} \right)^2 \end{equation*}$

This second-order form captures compressibility effects in the surface motion and is often sufficient for many practical applications in supersonic and hypersonic aeroelasticity.

For a thin surface of chord ${c}$ , the lift per unit span is obtained by integrating the pressure difference over the surface, i.e.,

(34) $\begin{equation*} L'(t) = \int_0^c (p - p_\infty) \, dx \end{equation*}$

Substituting the first-order piston theory approximation gives

(35) $\begin{equation*} L'(t) = \int_0^c \varrho_\infty a_\infty \left( \frac{2}{M_\infty} \frac{\partial w}{\partial t} \right) dx = \frac{2 \varrho_\infty a_\infty}{M_\infty} \int_0^c \frac{\partial w}{\partial t} \, dx \end{equation*}$

If the motion is harmonic, i.e.,

(36) $\begin{equation*} w(x,t) = \Re \left\{ \hat{w}(x) e^{i \omega t} \right\}, \qquad \frac{\partial w}{\partial t} = \Re \left\{ i \omega \hat{w}(x) e^{i \omega t} \right\} \end{equation*}$

then the lift becomes

(37) $\begin{equation*} L'(t) = \frac{2 \varrho_\infty a_\infty}{M_\infty} \frac{d}{dt} \int_0^c w(x,t) \, dx \end{equation*}$

For a pure plunging motion where the entire surface translates vertically as $h(t)$ , then $w(x,t) = \dot{h}(t)$ and so

(38) $\begin{equation*} \frac{d}{dt} \int_0^c w(x,t) \, dx = c \, \ddot{h}(t) \end{equation*}$

and therefore the lift per unit span becomes

(39) $\begin{equation*} L'(t) = \frac{2 \varrho_\infty a_\infty}{M_\infty} \, c \, \ddot{h}(t) \end{equation*}$

The corresponding sectional lift coefficient is defined as

(40) $\begin{equation*} C_L(t) = \frac{L'(t)}{\tfrac{1}{2} \varrho_\infty V_\infty^2 c} \end{equation*}$

so

(41) $\begin{equation*} C_L(t) = \frac{4}{M_\infty^3 a_\infty} \, \ddot{h}(t) \end{equation*}$

To express this result in a nondimensional form, assume harmonic motion $h(t) = h_0 \cos(\omega t)$ , so that $\ddot{h}(t) = -\omega^2 h_0 \cos(\omega t)$ , and introduce the reduced frequency $k = \omega c / (2 V_\infty)$ . Then the lift coefficient becomes

(42) $\begin{equation*} C_L(t) = -\frac{16 k^2}{M_\infty} \left( \frac{h_0}{c} \right) \cos(\omega t) \end{equation*}$

While piston theory provides a practical and computationally efficient tool for supersonic and hypersonic unsteady aerodynamics, it has significant limitations. It is valid only for small amplitude surface motions, assumes attached flow and quasi-one-dimensional wave propagation, and neglects shock waves, separation, and complex three-dimensional effects. The accuracy decreases significantly at low Mach numbers or near flow discontinuities. Despite these limitations, piston theory remains a foundational tool for high-speed unsteady flow modeling and is frequently employed in aeroelastic simulations and conceptual vehicle design.

Hypersonic Aircraft Design

A thin, flat plate is known theoretically to be the most efficient aerodynamic shape for hypersonic flight, its lift-to-drag ratio being the highest achievable. However, a plate is an unrealistic design solution for a practical flight vehicle because of the need to contain engines, fuel, systems, payload, etc. Wind tunnel measurements of candidate hypersonic shapes have shown that their maximum achievable aerodynamic efficiency in terms of lift-to-drag ratio, $L/D$ , decreases with increasing flight Mach number according to

(43) $\begin{equation*} \left( \frac{L}{D} \right)_{\rm max} = \frac{4 ( M_{\infty} + 3)}{M_{\infty}} \end{equation*}$

which is often referred to as Küchemann’s equation. The conclusions to be drawn here are that the aerodynamic efficiency is relatively low in hypersonic flight and that increasing the lift-to-drag ratio will be challenging.

To maximize the aerodynamic efficiency of a flight vehicle in hypersonic flight, it is recommended by experts in the field that some general design rules should always be followed. Such vehicles should generally be blended with body shapes with integrated forebodies, propulsion systems, and afterbodies. More specifically, they should have as small a frontal area as possible, with a highly streamlined overall shape to minimize surface area. A short-span, low aspect ratio wing should be used to keep the wing’s leading edge behind the Mach cone, with the fuselage suitably shaped to generate lift, i.e., as a lifting body. The propulsion system must be well-integrated into the vehicle’s overall shape to prevent shock waves from one component from adversely interfering with the flow at another component.

Waveriders

A waverider is another type of hypersonic wing design with an improved lift-to-drag ratio relative to other shapes. This goal is accomplished using the shock waves generated by its flight as a lifting surface, a phenomenon known as compression lift. The waverider concept originated from the work of Terence Nonweiler on winged atmospheric re-entry vehicles in the 1950s. A waverider is a delta wing with some anhedral, often called a caret wing. With the advent of computational fluid dynamics (CFD) and advanced hypersonic prediction methods, various optimized waverider shapes have been developed, as shown in the figure below. However, heat dissipation is an issue with such optimized designs because of the closer proximity of the shock waves to the surfaces.

A hypersonic waverider concept optimized to achieve the highest possible lift-to-drag ratio.

Propulsion Issues

Achieving combustion and propulsion is another major problem at hypersonic speeds, and the most likely candidate is a variation of a ramjet engine called a scramjet, or supersonic combustion ramjet. A scramjet, as shown in the schematic below, works on the principle that the shock waves on the vehicle can slow down and compress the air to support combustion inside the engine, without reducing the airflow to subsonic speeds. This contrasts with a conventional ramjet, where the incoming air is slowed to subsonic before combustion. In a scramjet, the flow remains supersonic throughout the entire engine cycle, avoiding the large total pressure losses associated with subsonic diffusion at high Mach numbers.

The scramjet’s principle relies on the vehicle’s motion compressing the incoming flow to support combustion.

Scramjets offer the possibility of efficient, sustained propulsion in the Mach 5 to 10 regime. Their ability to operate without onboard oxidizer and extract energy from atmospheric air makes them attractive for high-speed cruises and access-to-space missions. However, their dependence on high-speed entry, narrow operating envelope, and sensitivity to combustion stability and thermal limits remain open areas of research and development.

Combustion Process

The scramjet’s principle works on the Brayton cycle and relies on the vehicle’s forward motion compressing the incoming flow through a system of oblique and bow shocks. These shocks increase the air temperature and pressure to sufficient levels for combustion, even though the air remains moving at supersonic speeds inside the engine. Because of this, a scramjet cannot start or operate at low airspeeds, requiring the incoming air to be supersonic to ignite and sustain combustion. While it will operate inefficiently at lower Mach numbers, typically a scramjet must be accelerated to speeds near Mach 4 by another propulsion system, such as a rocket, turbojet, or ramjet.

Maintaining continuous combustion in supersonic flow is very challenging. The fuel–air mixture has an extremely short residence time in the combustor, typically just a few milliseconds, before it exits the engine. This short time makes it difficult to achieve efficient mixing and complete combustion. A rough estimate of the available combustion time is given by

(44) $\begin{equation*} \tau \ \propto \ \frac{L}{V} \end{equation*}$

where $L$ is the length of the combustion chamber and $V$ is the flow velocity, which remains supersonic. The challenge is further compounded by the formation of shock waves within the engine and complex shock–boundary layer interactions, which can lead to flow separation or instability that disrupts the combustion process. In addition, the high-speed flow produces severe aerodynamic heating, which requires the use of high-temperature materials and sometimes active cooling to manage the thermal loads on engine components.

Thrust Production

Momentum and pressure differences between the inlet and exhaust determine the thrust, $T$ , that a scramjet produces. A simplified thrust expression based on a control volume momentum balance is

(45) $\begin{equation*} T = \overbigdot{m}_e V_e - \overbigdot{m}_\infty V_\infty + (p_e - p_\infty) A_e \end{equation*}$

where $\overbigdot{m}_\infty$ is the incoming air mass flow rate, $\overbigdot{m}_e$ is the exhaust mass flow rate, $V_\infty$ is the freestream velocity, $V_e$ is the exhaust velocity, $p_e$ and $p_\infty$ are the exhaust and ambient pressures, and $A_e$ is the nozzle exit area. The fuel mass flow rate is denoted $\overbigdot{m}_f$ , so the total exhaust mass flow rate is

(46) $\begin{equation*} \overbigdot{m}_e = \overbigdot{m}_\infty + \overbigdot{m}_f \end{equation*}$

In the ideal case of perfect expansion, the exit static pressure matches the ambient pressure ( $p_e = p_\infty$ ), eliminating the pressure thrust term and reducing the thrust equation to

(47) $\begin{equation*} T = \left( \overbigdot{m}_\infty + \overbigdot{m}_f \right) V_e - \overbigdot{m}_\infty V_\infty \end{equation*}$

This form shows that thrust arises from the momentum gain of both the incoming air and the injected fuel. The added fuel mass, although small compared to the air mass, carries significant chemical energy, allowing the exhaust velocity $V_e$ to exceed $V_\infty$ . For good engine performance, the nozzle must be designed to produce $p_e = p_\infty$ at cruise Mach number. If the nozzle is not expanded correctly, then $V _e$ is reduced, and a pressure mismatch develops at the exit, degrading thrust and efficiency.

Specific Impulse

Unlike a rocket, which carries its oxidizer, a scramjet pulls in atmospheric air for combustion, making it more efficient in terms of specific impulse at high speeds. The specific impulse, defined as thrust per unit weight flow rate of fuel, is given by

(48) $\begin{equation*} I_{\text{sp}} = \frac{T}{\overbigdot{m}_f \, g_0} \end{equation*}$

where $g_0$ is the standard gravitational acceleration. Because a scramjet does not carry oxidizer, the total mass flow through the engine is much greater than the fuel flow alone, allowing for more momentum exchange with the atmosphere and greater thrust per unit of fuel consumed. This leads to significantly higher values of $I_{\text{sp}}$ than rockets, particularly in the Mach 5 to 10 range. Consequently, scramjets offer improved propulsive efficiency in high-speed atmospheric flight, although they are limited to operation within the atmosphere and require careful inlet and combustion control to maintain stability.

Thermodynamics

From a thermodynamic standpoint, combustion in a scramjet can be modeled as a high-speed, nearly constant-pressure process, similar to Rayleigh flow. The simplified energy equation is

(49) $\begin{equation*} h_\infty + q = \text{constant} \end{equation*}$

where $h_\infty$ is the stagnation enthalpy^[3] of the incoming flow, and $q$ is the heat addition per unit mass from fuel combustion. The flow gains total enthalpy from the chemical energy of the fuel, but experiences stagnation pressure losses because of shock interactions and viscous effects.

Although this framework describes the energy transfer in the flow, the amount of heat that can be added is ultimately limited by the thermal and structural capacity of the engine materials. The high-speed flow inside the scramjet produces intense aerodynamic heating, especially in the combustor and inlet regions. Surface temperatures can exceed 1,500 K, and may reach over 2,000 K depending on Mach number, altitude, and flight duration.

Material limits impose strong constraints on the maximum allowable wall temperature. Conventional aerospace alloys, such as titanium or nickel-based superalloys, are generally limited to about 1,000–1,200 K. Advanced ceramics, ultra-high-temperature composites, and refractory metals can tolerate higher temperatures, but are often brittle, expensive, or difficult to fabricate. Techniques such as regenerative cooling, thermal barrier coatings, or ablative surfaces may be employed to prevent thermal failure.

The maximum surface temperature a structure can reach in an adiabatic, uncooling case is approximately given by the recovery temperature, i.e.,

(50) $\begin{equation*} T_r = T_\infty \left( 1 + r \, \frac{\gamma - 1}{2} M_\infty^2 \right) \end{equation*}$

where $T_\infty$ is the freestream temperature, $M_\infty$ is the Mach number, and $r$ is the recovery factor, typically about 1.0 for laminar flow and 0.89 for turbulent flow. If this temperature exceeds material tolerances, active or passive cooling is required. Ultimately, the amount of energy that can be safely injected into the flow is limited not just by fluid dynamic choking effects but also by the capacity of engine walls to withstand thermal loads. These material constraints tightly couple thermodynamic performance to structural survivability in scramjet engine design.

Airframe/Engine Integration

One feature of a hypersonic airplane is the need for carefully integrating all of the parts of the aircraft that provide volume, i.e., for payload and fuel, with those for propulsion, i.e., an engine such as a scramjet. The figure below shows the basic idea for a “two-dimensional” configuration. An externally carried scramjet attached to the airframe (body) on a pylon would have an unacceptably high drag. However, properly integrating the body and the propulsion system could be expected to reduce the wave drag on the body and much of the external nacelle drag. The penalties paid for this achievement are a slight loss of useful volume and the ingestion into the intake of some of the boundary layers on the body surface. Studies have shown that careful aircraft/propulsion can be expected to increase maximum lift-drag ratios by about 1.5 times higher relative to a non-integrated design.

The principle of airframe/propulsion system integration for a hypersonic aircraft.

The Boeing X-51A scramjet demonstration aircraft was a waverider concept tested from 2011 to 2013, as shown in the artist’s impression below. The X-51A was used to demonstrate scramjet engine operations at hypersonic airspeeds and other aspects of hypersonic flight. The aircraft sustained Mach 5 flight for 200 seconds during its longest tests. Measurements made during these test flights have been used to compare measurements made in hypersonic wind tunnels and predictions made using CFD.

Image of the X-51A Waverider concept that is set to demonstrate hypersonic flight. — The X-51A Waverider was designed to demonstrate hypersonic flight. Powered by a scramjet engine, it could reach about Mach 6. (U.S. Air Force graphic.)

Sonic Booms in Hypersonic Flight

In hypersonic flight, a critical acoustic phenomenon is the sonic boom. As a flight vehicle travels significantly faster than the speed of sound, it generates intense shock waves that coalesce into a sharply defined, cone-shaped disturbance trailing behind the vehicle. These waves propagate through the atmosphere and, upon reaching the ground, are perceived as a loud and sudden boom. In the hypersonic regime, the increased shock strength and wave compression result in more concentrated and potentially more damaging sonic boom signatures. The pressure rise across the shock is much steeper, and the resulting wave fronts are sharper and more nonlinear than those in lower supersonic regimes.

Quantitatively, the peak overpressure $\Delta p$ of a sonic boom at ground level scales approximately linearly witht eh weight and with the square of the flight Mach number, i.e.,

(51) $\begin{equation*} \Delta p \ \propto \ W \, M^2 \end{equation*}$

This relation indicates that a vehicle flying hypersonically at Mach 6 can produce a ground-level boom several times stronger than one flying at Mach 2, assuming similar geometry, weight, and altitude. The high energy associated with hypersonic shocks also reduces atmospheric attenuation, allowing the boom to propagate further at the ground.

Mitigating sonic boom intensity is a significant challenge in hypersonic vehicle design, especially for applications involving flight over land. Strategies to reduce boom intensity include increasing flight altitude, streamlining the vehicle’s planform and nose shape to distribute shock strengths more gradually, and employing “low-boom” design principles, such as boom shaping, which involves tailoring the vehicle’s volume distribution to produce a pressure signature that deviates from the classic N-wave. Continued research into sonic booms at hypersonic speeds is essential for advancing high-speed flight, particularly supporting future commercial, military, and space access vehicles operating over populated areas.

Summary & Closure

Hypersonic flight vehicles operate in a regime where conventional aerodynamic assumptions no longer apply. At Mach numbers above 5, strong shock waves, high-temperature gas effects, and viscous–inviscid interactions dominate the flow field. The resulting kinetic heating can raise surface temperatures to thousands of degrees, imposing severe demands on materials and structural design. Hypersonic vehicles must incorporate specialized thermal protection systems and high-temperature materials, often involving ceramics, carbon–carbon composites, or ablative coatings to withstand these conditions. These constraints make thermal management a central aspect of vehicle design, tightly coupled to aerodynamic shaping and mission trajectory.

Propulsion remains one of the most challenging aspects of hypersonic flight. While scramjets (supersonic combustion ramjets) offer a promising solution for air-breathing propulsion in this regime, their operational stability, ignition, and integration with the airframe remain active areas of research. One of the most promising aerodynamic configurations is the waverider, which generates lift by aligning its lower surface with the shock wave it produces during flight. By effectively “riding” its own shock, the waverider minimizes wave drag while maximizing lift-to-drag ratio, offering improved efficiency at hypersonic speeds. The future of hypersonic flight will likely depend on continued advances in propulsion, high-temperature materials, and integrated vehicle design, culminating in systems capable of sustained, efficient, and potentially reusable high-speed atmospheric flight.

5-Question Self-Assessment Quickquiz

For Further Thought or Discussion

What defines hypersonic speed, and why is it different from just “very fast”?
Why does aerodynamic heating become a significant issue at hypersonic speeds?
What kinds of shapes work best for hypersonic vehicles, and why?
Why can’t regular jet engines work at hypersonic speeds?
How do shock waves form, and why are they stronger at higher Mach numbers?
What are some challenges in keeping a hypersonic vehicle stable in flight?
Why do hypersonic vehicles fly high in the atmosphere?
What kinds of materials are needed to survive the extreme heat of hypersonic flight?
Why is testing hypersonic vehicles on the ground so difficult?
How might hypersonic vehicles be used in the future, e.g., for space, military, or travel?

Other Useful Online Resources

For additional resources on hypersonic flight, follow up on some of these online resources:

Hypersonic Waverider – How the USAF X-51A scramjet works – see the video here.
Hypersonic Aerodynamics: Basic and Applied – start to watch the multi-part lecture here.
2025 AIAA Durand Lecture by Kevin Bowcutt – view the talk here.
Testing and Analytical Challenges in Hypersonics – Dr. Mark Lewis – video available here.
Hypersonic Flight Vehicle Design and Performance Analysis (AIAA short course) – see course info here.
Hypersonic and High-Temperature Gas Dynamics by John D. Anderson – PDF available here.
The Insane Engineering of Re-Entry – video available here.

Entropy is a thermodynamic quantity that measures a system's disorder or randomness. ↵
In aerodynamic flows, entropy increases when energy is degraded by shocks or viscous dissipation. ↵
Enthalpy is a thermodynamic property that represents the total heat content of a system. ↵

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Introduction to Aerospace Flight Vehicles Copyright © 2022–2025 by J. Gordon Leishman is licensed under a Creative Commons Attribution-NonCommercial-NoDerivatives 4.0 International License, except where otherwise noted.